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Device for separation and opening of aircraft solar battery flaps Invention relates to space engineering and may be used in deploying solar batteries of spacecrafts. Device for separation and opening of aircraft solar battery flaps (DSOSBF) includes frame, two packs of flaps, hold-down locks with hooks, rockers, spring-loaded tie rod with yokes interacting with upper flaps, spring-loaded rods with through holes for pin with butt ends interacting with profiled grooves, retaining rings with retainers. |
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Rotation system of double-wing solar panels Invention relates to orientation (rotation) gear of the solar panels. The rotation system of the solar panels contains casing (1) with covers (2), output shaft in form of two parts (3) and (4) with flanges (5) and (6) to connect with the solar panel wings, and central drive (7). Free from play gear drives (8), (9) connect parts (3) and (4) of the output shaft via the drive (7) shaft. Theses gear drives exclude play and increase torsional rigidity of the output shaft. To exclude play in bearings (10), (11) of supports of the output shaft parts the Bellevible springs (12) are provided. On surface of parts (3, 4) the power current collectors (13), (14) of solar panel wings are installed, and inside the remote current collectors (15), (16) are installed. PCBs (17, 18) of the drive (7) are the output shaft supports (3, 4) ensuring increased bend strength of the structure. |
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Method for orientation of artificial earth satellite 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). |
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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. |
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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. |
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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. |
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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. |
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Method of controlling orbiting spacecraft 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. |
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Navigation satellite orientation system 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. |
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Control over orbital spacecraft 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. |
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Method for generation of control actions on spacecraft 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 2 R . Number of flows is n = ( S x / 2 R ) − 1 . By mutual bias of flows in direction of their motion for 2 R distance droplet mist flows are generated in number of m = ( S y / 2 R ) − 1 . Each of the mentioned flows is biased relative to previous flow for 2 R 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. |
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Solar battery for small-size spacecrafts and method of its manufacturing 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. |
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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. |
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Invention relates to spacecraft electric power supply with the help of solar batteries. Proposed method comprises definition of preset and current angles of solar battery orientation and solar battery angular velocity (ωSB). Design angle is computed to assign measured angle magnitude thereto and memorised prior to start of control over solar battery. Solar battery is spinned in direction of decrease in mismatch between preset and design angles. Defined are angles of solar battery acceleration and deceleration (tAC, αDEC) and threshold (tTHR, αTHR) and maximum tolerable angle of its deflection (αMAX) proceeding from minimum tolerable currents of solar battery. Said angles are used to set operation threshold (αS). The latter exceeded, said mismatch is generated. The latter is not taken into account if lower than drop-away threshold (αDROP). The latter reached, solar battery spinning is terminated. Solar battery design angle is corrected with the limits of one discrete sector of solar battery spinning circle. Discrete sector magnitude depends of angles αAC, αTHR and αS. Depending upon αS and ωSB threshold of the interval of control over continuous variation of data on solar battery angular position is set. Count of said interval is made if current measured angle differs from memorised one by more than one discrete sector and is terminated otherwise. Threshold of the time of control over solar battery spinning is set depending upon tAC, tTHR, αMAX, ωSB and discrete sector magnitude. This time is counted at zero time of control over continuity is sign of mismatch between measured and memorised angles dose not satisfy the solar battery preset direction of spinning. Otherwise, count is terminated to zero the time of control of spinning direction. Note here that when measured angle varies by one discrete sector, angular angle of boundary between discrete sectors is taken to be design angle to assign new measured angle to memorised angle. In case the time of control over continuity or that over spinning direction exceeds its threshold, failure signal is generated to terminate control over solar battery. |
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Invention relates to spacecraft electric power supply with the help of solar batteries. Proposed method consists in definition of preset angle of solar battery, measurement of its current angle and computation of design angle by angular velocity and spinning time. Solar battery acceleration angle (αAC) and deceleration angle (αDEC) are defined. Solar battery is spinned to threshold of drop-away (αDROP ≈ (αDEC) when mismatch between said preset and design angles is terminated. Before start of the control, preset angle is memorised to take initial preset angle as valid actual current angle. Mismatch threshold (αTHR) of said angles proceeding from angles αAC and αDEC, as well as minimum tolerable currents of solar battery. Angle transducer circle is divided into equal discrete sectors by magnitude a given the condition αAC + αDEC < σ < αTHR. Discrete sector bisectors are taken to be measured magnitudes. Period of valid current angle definition is set to the order and exceeding maximum duration of transducer data fault and smaller than minimum interval of faults train. Said interval is divided to four equal interval while analysis of measured and memorised magnitudes in said intervals are reset to generate validity signal. In the latter case, solar battery is spinned to mismatch between design and preset angles αDROP to set new preset angle magnitude. |
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Solar battery strut comprises two-link mechanism with common axle supporting torsion spring with cocking devices. One link is arranged at solar battery frame while another one is mounted at craft body. Spring-loaded rod to lock the link at end position is arranged at said link perpendicular to axis. Rocker is arranged at spring-loaded rod end to turn thereat. Antifriction bearings are rigidly secured at rod both ends to interact with cam taper grooves, said cams being rigidly mounted at the link opposite spring-loaded rod. Links of aforesaid mechanism have openings to link retainers threaded therein. |
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Space solar electric station and independent photo emitting panel Set of intentions relates to space power engineering and may be used for transmission of electric power in the form of laser radiation to Earth surface and for high-accuracy measurements in space, data transfer, etc. Proposed station comprises base module 1, system of mirrors 2, laser radiation summator 3 directed to system 2 and photo converter panel 4 arranged outside of module 1. Every panel 4 consists of two types: photoelectric panels 5 and independent photo emitting panels 6. The latter are connected in chain for self-opening and arranging in closed flat zigzag-like figure. Panels 5 are mounted at the start of chain 5, 6. Note here that the first panel is connected with base module 1. Said module 1 comprises the following systems: control system 8, cooling system 11 and supply system 12. Every panel 5 is connected with supply system 12. Every independent panel 6 is composed of a carcass with Fresnel lenses are carcass end with photo converters (not shown) aligned therewith and located there above. Carcass bottom part base accommodates power accumulators, control unit of panel 6 and fiber lasers with pumping units and laser radiation summator. Aforesaid photo converters of panel 6 are electrically connected via power accumulators with pumping and control units. Summators of independent panels 6 are connected to aforesaid summator via FO 30. |
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Invention relates to space engineering and may be used in designing external structures of spacecraft, primarily, solar cell batteries. Solar cell battery comprises frame, articulated top and bottom flaps with torsions fitted on hinge pins. Opposite ends of torsions support brackets wherein fitted are torsion resetting mechanisms. Said brackets are secured at torsions and set to initial position, in symmetry about torsion axis. Note here that one of said brackets is fitted on top flap while second bracket is mounted at bottom flap to allow resetting mechanism to twist torsions in one direction. |
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Bench for opening panels of solar battery Invention relates to ground tests of opening structures, predominantly solar batteries (SB) with null-gravity conditions simulation. The bench is designed for opening two dissymmetrical SB panels (1) and contains frame leg (2) on which weight-releasing device (3) is mounted and adapter frame (4) for spacecraft simulator (5). In the upper part of leg (2), bracket (6) is installed. The bracket is moved horizontally. In the lower part of leg, adjustable pillars (8) are installed. Device (3) is made as separate swivel links (9) where bracket (6) is rigidly connected with the first link. Rotation axes of links are coaxial to rotation axes of corresponding SB panels (1). The latter is provided by moving the bracket (6) manually along guides and by fixing it with special screw. In each link (9), two dampers in the form of rods (not shown) are fixed. During SB testing for opening connection with board (5) is released and SB (1) flaps begin to open under action of operational springs. As gap between SB (1) flaps and device (3) is limited (not more than 150 mm) the presence of the said dampers with hangers has little altering influence on calculated flap movement, and their rigid connection in the form of rods provides synchronous movements of SB (1) flaps and the device (3). |
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Proposed system comprises casing, hollow shaft with solar battery connection flange, solar battery rotation drive, power and telemetry current collection devices. Output shaft is made up of structural flange and shaft with power current collection device. Telemetry current conducting device is fitted on its shaft and engaged with output shaft. Output shaft flange is arranged in solar battery turning system casing to run in thrust bearing either with preload or pressed via thrust bearing against said casing. |
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Method of spacecraft solar battery position control during partial failures of aspect sensor Invention relates to electric power supply systems of spacecraft (SC). According to the method, SC solar battery (SB) is rotated at sustained design angular velocity by an order or more greater than angular velocity of SC circulation on orbit around the Earth. Angular position of direction to the Sun unit vector projection on normal to SB working surface rotation plane in coupled coordinates. Full circle of aspect sensor is split into equal discrete sectors each one of which has corresponding arbitrary value at sensor output. Preset angle increment divisible by discrete sector dimension is set. Preset angle is calculated as integer number of this angle increments in angular position of specified projection of normal unit vector. Thresholds of operation and release as well as initial SB angular position where the mentioned normal coincides with bisecting line of one of angular sectors are set. Initial value of design angle (as product of design angular velocity by rotation time) corresponds to angular position of this normal. At the moment when sensor readings change, design angle is assigned value equal to number of discrete sectors contained in it increased by half of sector. Threshold value of control time is set exceeding time of SB rotation by angle with maximum number of angular sectors with equal output values. During SB rotation, time of design angle correction is counted. Failure signal is produced if correction time reaches or exceeds threshold value of control time. |
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Method of spacecraft solar battery orientation by electric current Invention relates to electric power supply systems of spacecraft. The method includes presetting the solar battery (SB) design angular rotation velocity exceeding the angular velocity of aircraft circularisation around the Earth by an order and more. In this process, electric current generated by SB is measured, SB rotation period is determined from mismatch between its design and preset angles down in this mismatch. SB is rotated prior to control start up to the moment of starting to reduce achieved greatest amount by means of current. The design SB angle is assigned the value of preset angle, herewith, the design angle is calculated during SB rotation as product of design angular velocity by SB rotation time. Time points of maximum current generation between moments of rotation termination and starting the next rotation are stored in memory. The design angle is corrected for amount of: error correction depending on aircraft angular orbital velocity, stored in memory time values at the moment of SB rotation termination and at the moment when current generated by SB reaches maximum value, as well as on SB rotation period. Design angle is increased or decreased by error correction value when SB is rotation during time reckoning in direction of increasing or decreasing angle respectively. |
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Device for electromechanical linkage between fixed structural parts of spacecraft Invention relates to space engineering, particularly, to spacecraft fixed structural elements electrically connected with spacecraft control system, for example, solar batteries, antennas, moving covers, etc. Proposed device represents loop cable 7 secured by three yokes to solar battery moving part, solar battery drive and rigidly secured on spacecraft body. Yoke 10 divides loop cable 7 into two parts and is rigidly secured to axle 11 fitted in fork 12. Said fork 12 is pivoted to drive support 13 of turn drive. With fork 12 turning in opening the solar battery sections, roll displacement of the section of loop 7 arranged between said first and second yokes is tracked. Yoke 10 revolves about axis perpendicular to form turn axis. |
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Solar cell panel for spacecraft Invention relates to spacecraft electric power supply systems and specifically to solar cell panels. The solar cell panel comprises two panels, each consisting of two half-panels, having pivotally connected and series-assembled into a bundle a root (2), middle (3) and an outermost (4) flap. The flaps are mounted on a frame (5) which is movably mounted on four mounting assemblies (6) of the body (1) of the spacecraft. The half-panels are connected to each other on one side by four spring-loaded clamps (7), and on the other by four couplers (8) in the mounting assemblies (6). The root flap is connected to the middle flap by an axle (9) and the middle flap is connected to the outer most flap by an axle (10). The spring-loaded clamps are connected by a rope to a pyro device (not shown). Two supporting arms are mounted in pairs on each flap. The supporting arms mounted on the outermost flap are provided with axles which interact, when opening the panels, with profiled protrusions on the supporting arms mounted on the root flap. This enables to open the panels using a "roll" technique, wherein there is organised retraction of flaps, which prevents their collision with spacecraft equipment when opening. The number of pyro devices in locking elements of the panels is reduced. |
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Invention relates to control of orientation of space vehicle (SV) with immovable relative to SV body panels of solar batteries (SB). Control method includes SV gravitational orientation and its spin around longitudinal axis (minimum momentum of inertia). When the Sun is near orbit plane this plane is aligned with SB plane by the time of passing morning terminator. Angle between normal to SB active surface and direction to the Sun is measured and monitored. At the moment of passing morning terminator, SV spin is performed in direction corresponding to decrease of the mentioned angle, where spin angular velocity is selected from the range of 360°/T - 720°/T, where T is SA orbiting period. |
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Unit offices and opening the valves of solar batteries of spacecraft Invention relates to a drop-down designs of spacecraft such as solar cells (SC) or an antenna. The device consists of a frame rigidly fixed to the shaft of the actuator, and two packages of wings, lower wings Packages (3) fixed to the frame freezes and the average leaf (4) are connected with the lower sash (3) and the upper hinge leaf. In the axis of hinges mounted cocked springs (torsion), revealing the leaf into position. The frame and sash packages with pyro devices attached to the body of the spacecraft. Package leaflets along the frame rigidly fixed clamping locks. The main body of clamping lock hinged hooks (9), interacting with spring rods (10) of these locks. Hooks (9) also interact with the rockers, which are enshrined in the spring-powered (held under the hooks (9), and rockers). Explosive pins are rigidly fixed on the frame interacting with turning a lever attached to said pivotally drawn. When submitting the team for at least one of the couples latest release spring-loaded rod from the cocked position by rotating said lever. In the case of activation of the two spring-loaded rod, explosive pin moves without rotation of the lever. When you move the spring-loaded rod turned all rocking along the package leaflets. Hooks (9) of all the action clamping locks the spring-loaded rod (10) is also rotated, ensuring the free movement of these rods and valves opening solar batteries. |
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Device to illuminate photoelectric converters of spaceship solar battery Invention relates to ground equipment for servicing spaceship with solar batteries. Proposed device comprises corrugated jacket 9 made of lightproof material and arranged on case 1. Said case houses pulse electric radiators 6 arranged in cells 7 formed by intersecting ribs 8. Frame is secured on end face of aforesaid corrugated jacket 9 to interact with solar battery section carcass along said carcass (not shown). Case 1 is equipped with height-adjustable support posts while light reflecting coat is applied onto inner surface 13. Fans may be arranged inside said case. Device may be used in check switching on of spaceship onboard hardware at space center complex. It allows testing serviceability (illumination) of photoelectric converters of section solar batteries and integrity of circuits that pick up electric power therefrom. It ensures preventing intolerable overheat of optical pulse electric emitters (LEDs). |
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Invention relates to aerospace engineering, namely, to devices intended for direct conversion of solar energy into electric power by means of photo cells. Proposed panel comprises carcass with transverse sections, photo cells modules and load-bearing surface made up of elastic elements that include lengthwise strings connected with carcass Said load-bearing surface comprises additionally crosswise strings also connected to carcass via bushes fitted in carcass taper holes to adjust tensioning and locking strings in bush bore. Transverse strings are arranged above lengthwise strings secured in dielectric bosses fitted on carcass crosswise sections. |
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Spacescraft power supply control method and system to this end Invention relates to spacecraft power supply. Proposed method involves replenishing spacecraft power supply from external sources. One or several spacecraft-mounted electric power stations are located on working orbits in spacecraft line of sight. Spacecraft location is determined to cut on spacecraft tracking system and electromagnetic power is transmitted to spacecraft onboard receiver. Transmission can be performed in the range of laser to microwave radio radiation, or in the form of high-power electron beams. Spacecraft power supply system in normal state, supply of electromagnetic power from spacecraft-mounted electric power station is terminated and it is moved to standby orbit. Proposed system comprises device to transmit electric power arranged on spacecraft-mounted electric power station platform driven by rocket engine. Said platform carries also laser range finder optically connected with one or several angular reflectors arranged on spacecraft. Said reflectors are used to align spacecraft-mounted electric power station conducting channels with spacecraft electric power receiving channels. |
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Method of determining ground albedo in subsatellite points of spacecraft orbit (versions) Invention relates to space engineering, used in geophysics for determining and controlling integral parametres of radiant heat exchange of sections of the surface of a planet, around which spacecraft with solar cells (SC) is rotating. The method involves determination of moments when the sun is in a zenith region above the said spacecraft (and the spacecraft, respectively - in the point under the sun). If the spacecraft is fitted with two single-sided solar cells with zero power output of their back surface, then when one solar element passes the said point under the sun, then one solar cell is turned normally in the direction of the sun, and the other solar cell is turned in the direction opposite the sun. Value of current in each solar cell is measured, from which, taking into account a known coefficient of relative output power of the battery, the albedo of the corresponding subsatellite point of the earth surface is determined. For spacecraft, the solar cells of which have back surface with non-zero output power, at the moment of passing through points under the sun on two successive orbit passes, one ore more solar cells are turned normally in the direction of the sun on the first pass. At the second pass, turning is done in an opposite direction. Values of current of the solar cells in their first and second orientations are measured, from which, taking into account the known coefficient of relative output power of the back and working surfaces of the solar cell, the albedo in the corresponding subsatellite point of the surface of the earth is determined. |
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Method of controlling space vehicle solar battery position Measurement of present angular position of solar battery (SB) and determination of the given Sun orientation. In the event of difference between the given direction and SB current angular position the SB rotation instructions are formed and rotational speed of output shaft of SB rotation device is set up. This speed is a sequence higher than rotational speed of spacecraft around the Earth. Before the control has started time and SB deceleration angle is determined from the instruction issue to the full stop. Angle threshold value is determined as angle of less than 180º and exceeding the sum of deceleration angle and angular value of the same discrete sector of SB angle transmitter. The time control threshold value is set as the ratio of angle control threshold value to the SB rotational speed. After rotation cessation instruction has been formed, the current SB position is stored and running control time is measured. The current control time measurement is ceased and its value is set to zero if SB clockwise or anticlockwise instruction has been issued before the current control time gains threshold value. At the moment of gaining the threshold value by current control time the current angular SB position is measured and compared with its stored value. If the current angular SB position differs from the restored one by value exceeding the control angle threshold value, SB control is stopped and a failure signal is initiated. |
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Low-thrust electro-thermal rocket engine Electro-thermal rocket engine (1) heating compartment (20) for fluid working medium (4) to pass there through before being ejected from nozzle (8). Working medium is heated by either ohmic or electric-arc method. Note also that spacecraft comprises electric power source (22) incorporating photo galvanic elements (24) fitted on heat exchanger (10). Working fluid (4) passes through heat exchanger upstream of aforesaid heating unit (20). Electric power source (22) comprises concentrator (28) directing sun light onto photo galvanic elements (24). Heat generated on the latter is transferred to working medium (4). |
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Method of controlling space vehicle solar battery position Measurement of present angular position of solar battery (SB) and in the event of difference between it and given direction, SB rotation instruction is formed. In the event of no difference, SB rotation cessation instructions are formed and constant sufficiently high rotational speed of output shaft of SB rotation device is set up. Time and SB boost angle is determined from rotation instruction issue to steady-state value of SB rotational speed. The same is determined for SB slowdown after rotation cessation instruction has been issued. The drop-away of less than 45° and not less than 0° is set. Operation threshold of not less than the sum of acceleration and release angles and not exceeding 45° is set. Error signal is formed, if the angle between the specified angle and present one exceeds the operation threshold. This signal generation is stopped, if during SB rotation error angle will be less than release threshold or equal to 0°. The same is done, if error angle character at the beginning of rotation does not coincide with error character at the end of rotation where the specified angle does not exceed the operation threshold. Control time threshold value is set as the ratio discrete sector of angle transmitter to SB sustained rotation rate during acceleration and deceleration. SB rotation time is controlled up to rotation cessation instruction. At the end of deceleration the current control time as well as moments of rotation redirection shall be reset. Rotation cessation instruction and failure signal for SB rotation controlling device is formed when current control time exceeds control time threshold value. |
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Method of controlling space vehicle sun battery position Proposed invention relates to systems of supplying space vehicles (SV) and can be used aboard various geostationary satellites. The proposed method comprises measuring the current angle of sub battery (SB) that defines the position of normal to SB operating surface in coupled coordinate axes. In control, SB is revolved relative to SV housing, opposite the direction the said housing revolves relative to the Sun. SV complete geostationary orbit is determined. Note here that zero time is taken to be the moment at which aforesaid normal to SB surface coincides with projection of unit vector of direction to the Sun onto the said normal plane of rotation at zero current angle of SB. Current orbital time is count from zero to SV period around the Earth. Note also that the SB rotation circle is divided into equal arcs proceeding from maximum and minimum tolerable currents generated by SB at its certain orientation relative to the Sun. Bisectrix of SB rotation initial angular segment is taken to make the SB current angle zero value. SV complete period around the Earth is divided into time intervals, the number being equal to the number of the said segments. Each time interval is brought into correspondence with the current angle at a certain boundary between angular segments. SB rotation is terminated every time the aforesaid segment coincides with the boundary between angular segments assigned to SB. |
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Method of determining spacecraft solar battery maximum output Propose method designed to determine spacecraft solar battery maximum output comprises measuring the spacecraft orbit altitude for the angular half-width of the Earth disk visible from the spacecraft (Qz) and the angle of elevation of atmosphere top boundary (ε) above the Earth horizon visible from the spacecraft, determining the angular half-width of the Sun disk visible from the spacecraft (Qs) and measuring spacecraft orbital angular velocity (ω). It includes also measuring the angle between direction to the Sun and spacecraft orbit plane (β), measuring the angle of elevation of direction to the Sun above the Earth horizon visible from spacecraft (g) and maximum solar battery output at the batteries minimum temperature. The system, to this end, comprises the unit to measure the angle between direction to the Sun and spacecraft orbit plane, unit to measure the angle of elevation of direction to the Sun above the Earth horizon visible from spacecraft, the unit to measure spacecraft orbital angular velocity, the unit to measure the spacecraft orbit altitude variation, the unit to determine the angle of elevation of atmosphere top boundary (ε) above the Earth horizon visible from the spacecraft, the unit to determine the angular half-width of the Sun disk visible from the spacecraft. It incorporates also the unit to determine the orbit whereon solar battery maximum output is determined, the unit to generate identification parameters of solar batteries temperature conditions and key. |
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Method of determining spacecraft solar battery maximum output System designed to determine spacecraft solar battery maximum output comprises a unit to measure the angle between direction to the Sun and spacecraft orbit plane, unit to measure the spacecraft altitude, unit to measure spacecraft angular velocity, unit to define the orbit enlightened section beginning, unit to determine time intervals allotted for determination of solar battery maximum output, time oscillator, comparator unit and key. |
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Invention refers to space equipment, to space vehicle electric power supply systems, and can be used during solar battery operation. Method for determining maximum power output of space vehicle solar batteries involves measurement of angle between the Sun direction and space vehicle orbit plane, the Sun direction elevation angle above the Earth horizon observed out of space vehicle, and maximum power output of double-sided solar batteries and solar batteries having positive power output of their back surface, which is defined as product of voltage and current values from solar batteries, which are measured at the moments when radiation reflected from the Earth enters solar battery panels from their end surface, which are defined from condition of equality of values of the Sun direction elevation angle above the Earth horizon observed out of space vehicle and one-half angle of sensitive area of solar battery panel operating surface. Maximum power output of single-sided solar batteries is defined as product of voltage and current values from solar batteries, which are measured at the moments when radiation reflected from the Earth enters solar battery panels from their end or back surfaces. System for determining maximum power output of solar batteries of space vehicle includes measuring unit of angle between the Sun direction and space vehicle orbit plane, unit for determining the loop of executing the determining operation of maximum power output of solar batteries, for measuring the Sun direction elevation angle above the Earth horizon observed out of space vehicle, unit for determining the moment of determination of maximum solar battery power output, and a key. |
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Method for orientation of satellite solar battery Method includes superposition of normal line to illuminated working surface of solar battery (SB) panels to plane created by axis of panels and direction at the Sun. Angle between direction to the Sun and specified normal is established and maintained so that value of SB power (measured as product of its current by voltage) corresponds to daily consumption of the Earth satellites loads. This angle is also used to provide maintenance of SB voltage at required level in the point of its current-voltage characteristic that corresponds to maximum output power of SB (at preset angle of SB establishment). In case specified voltage of SB reduces at accurate orientation of SB at the Sun, SB is inverted from direction to the Sun by the angle, at which specified voltage of SB remains at required level. |
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System of control of spacecraft solar batteries position System contains solar batteries (SB) turning device, amplifying-transforming device, SB solar-orientation control block, SB preset position alignment block, current controller block, current sensor, power system control block. System additionally includes devices for measuring: spacecraft (SC) orbit altitude measuring, SC orientation, Sun elevation angle over Earth horizon visible from SC. There is also block assigning maximum value of current produced by SB under influence of direct solar radiation and blocks determining: moments of ingress of radiation reflected from Earth on rear side of SB, moments of generation by SB of additional power under influence of radiation reflected from Earth, turning angle of SB, and area of SB work surface lit by solar radiation. The circuit also includes two keys and elements "NOT" and "OR". |
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Method of cosmic vehicle solar batteries position control and system for its implementation Method includes turning of SB at pre-set angle between normal to SB work surface and sun direction. At that the angle between Sun direction and cosmic vessel (CV) orbital plane is measured. At moments when completion of 90° angle of visible half-spread of Earth exceeds the said measured angle and on condition CV is located at Sun-lit part of orbit, the angle between Sun direction and local vertical and Sun elevation angle over visible from CV Earth horizon are measured. In case the said elevation angle exceeds 90°, the moments of penetration of the Sun radiation reflected from Earth on back side of SB panels are determined. The current produced by SB is measured and if it exceeds value of current produced by SB with work surface perpendicular and if the sunrays reflected from Earth fail to reach, SB are turned away from Sun direction. Turning away is performed until position is reached at which the current produced by SB from direct Sun radiation on their work surface and the one produced by the sunrays reflected from Earth on back side of SB panels reach maximum possible value at the current position of CV on orbit. The suggested control system includes units and interconnections between them necessary for performance of the above-described operations. |
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Method of spacecraft sun batteries position control and system of its implementation Inventions are related to power supply of spacecrafts (SC) by means of sun batteries (SB). Suggested method includes turn of SB panels into working position that corresponds to alignment of perpendicular to their illuminated working surface with the plane that is formed by SB panels rotation axis and direction to the Sun, and turning of SB panels at the set angle between specified perpendicular and direction to the Sun. The angle of the Sun elevation above the local horizon is measured on the illuminated part of SC orbit and is compared to the value of angular SB sensitivity area semi-opening. In the moments when this angle is exceeded by this value, present value of SB current is measured and compared to the current that is generated by SB, when their working surface is oriented to the Sun, and there is no light reflecting from the Earth hitting the SB surface. When value of present current from SB is exceeding value compared to it, SB are turned from direction to the Sun to the side of the Earth centre at a certain angle so that light flows, direct and reflected from the Earth, are hitting SB. Suggested control system includes the necessary units and connections between them for performance of above mentioned operations. And it also includes unit of SC illumination moments determination, unit of SC orbit height, unit of angle measurement between directions to the Sun and horizon, unit of setting value of current from SB from direct sun radiation, unit of determination of time points of SB maximum current exceeding set value, unit of moments determination of reflected sun radiation hitting SB working surface, unit of SB turn control, and element "И". |
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Method of spacecraft sun batteries position control and system for its implementation Inventions are related to power supply of spacecrafts (SC). Suggested method includes turning of sun batteries (SB) panels into working position, when the perpendicular to the illuminated SB surface is aligned with the plane that is formed by SB rotation axis and direction to the Sun. At that commencement of approach floe negative effect to the working surface of SB is determined; SB panels are turned until this period of time and SB are returned into the working position on completion of this effect. In the moments of SC illumination by the Sun, the angle is measured between direction to the Sun and direction of SB panels' rotation axis. In the moments when load current value is exceeded by the maximum possible current from SB, SC speed vector direction is measured and area of SB panels projection to the plane that is perpendicular to this vector is measured. If this area exceeds its threshold value, SB panels are turned at the angle that corresponds to the minimum effect of approach flow, with simultaneous provision of SC with power supply. Whenever SC is located in the Earth's shadow, SB panels are turned so that their plane is aligned with SC speed vector. Suggested system contains drives and units that are necessary for SB position control, and it also includes unit of SC illumination moments determination, unit of angle measurement between direction to the Sun and direction of SB turning axis, unit of determination of time points of SB maximum current exceeding load current. It also stipulates for unit of SC speed vector direction measurement, unit of area determination of SB panels projection to the plane that is perpendicular to SC speed vector, and unit of SC orientation control along with SC speed vector direction. |
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Invention is related to the area of spacecraft (SC) control. Suggested method includes setting the viewing axis of probing instrument (PI) relative to construction lines of SC and SC rotation until PI viewing axis is aligned with direction to the Sun. Also SC is rotated until axis of its minimum moment of inertia is aligned with orbit plane. Height of SC orbit is measured, defining the angle values between direction to the Earth's centre and directions to the lower (λ0) and upper borders of investigated atmosphere layer. Depending on this angles, PI viewing axis is set at certain angle (λ) to the axis of SC minimum moment of inertia to the side that corresponds to the highest illumination of SC sun batteries. This angle λ accounts for minimum mismatch between the current angle of SC minimum moment of inertia axis deviation from the local vertical line and angles of this axis deviation from local vertical line in the beginning and in the end of probing session. The angle is measured between direction to the Sun and SC orbit plane, and if it agrees with value of angle λ0, then angle of the Sun elevation above the Earth's horizon visible from the SC. If the values of the latter angle are less or equal to angle of investigated atmosphere layer upper border elevation above the Earth's horizon visible from the SC, SC is turned until PI viewing axis is aligned with direction to the Sun. At that axis of SC minimum moment of inertia is aligned with the direction that lies in the orbit plane and forms a certain angle with projection of direction to the Sun to this plane, and this angle depends on above mentioned angles λ0 and λ. Atmosphere probing is performed in the moments of the Sun setting behind the Earth's horizon visible from the SC, maintaining the constant orientation of the SC sequentially in orbital and inertial coordinated systems. |
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Inventions are related to the area of spacecraft (SC) control. According to the suggested method, SC is stabilized and Earth's atmosphere is probed in the moments of the Sun setting behind the Earth horizon visible from the SC. At that SC is rotated by axis of its minimum moment of inertia perpendicularly to the orbit plane. Height of SC orbit is measured and probing instrument (PI) viewing axis is set to the side that corresponds to the highest electric power collection from the SC sun batteries, for a certain angular distance from the axis of SC minimum moment of inertia. The angle is measured between direction to the Sun and SC orbit plane, and this angle is compared to the angle between the direction to the lower border of investigated atmosphere layer and direction to the Earth's centre. In case they coincide and if the angle between axis of SC minimum moment of inertia and perpendicular to orbit plane is minimum, SC is turned until PI viewing axis is aligned with direction to the Sun. Angle of the Sun elevation above specified visible horizon is measured and this angle is compared to the angle value of elevation above the upper border horizon of the Earth's atmosphere investigated layer. In the period, when the angle of the Sun elevation is less or equal to the value compared to it, the Earth's atmosphere probing is being done. The suggested control system includes necessary units and connections between them for performance of above mentioned operations. And it also has incorporated measurement units for above mentioned angles measurement and estimation of atmosphere probing moments and development of according orientation of SC. |
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Method of control of spacecraft solar battery position and system for realization of this method 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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Method of control of spacecraft solar battery position and system for realization of this method 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. |
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Method of control of spacecraft solar battery position and system for realization of this method 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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Proposed method includes placing the flexible sectors on carrier, rotating the carrier in plane corresponding to working position of frameless centrifugal structure and deploying the sectors from carrier under action of centrifugal forces. Sectors are interconnected by side edges forming single working surface in the course of their deployment and preliminarily when necessary. Additional deploying force is applied to sectors along joint areas from periphery of centrifugal frameless structure to its center. Device proposed for realization of this method has carrier for placing the sectors, its rotation drive, as well as drive and mechanism for extension of sectors. Articulated on bearing part of extension mechanism are brackets provided with pairs of hold-down and drive rollers. One sector joint area is passed through each pair of roller. Drive roller is provided with drive which is kinematically aligned with sector extension drive, thus forming additional deploying force. Side edges of sectors may be connected at points of application of this force (or near them) with the aid of connecting elements, such as zipper, as well as by welding, bonding, sewing, etc. Carrier may be made in form of common drum or in form of reels separate for each sector. Device forming the centrifugal frameless structure has surface smoothly stretched in two axes. |
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Spacecraft solar battery of large area Proposed solar battery has central power member. Solar battery consists of two sections. Each section is formed from standard film-type trihedral prisms on base of inflatable tubular skeleton. Outer surface of this skeleton is coated with compound which gets hardened under action of ultraviolet and visible solar radiations. Solar battery deployment system includes two electric motors of central power member and additional electric motor. Inputs of these electric motors are connected with outputs of pitch, yaw and roll channels of solar battery control unit. Solar battery is additionally provided with additional position electric motors which are used for discrete turn of each trihedral film-type prism through angle of 0o to 360o at pitch of 120o. Specification gives description of solar battery modification which includes reserve film-type panels increasing active life of solar battery. Total power of proposed solar battery is about 120 kW. |
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Method of control of spacecraft power supply system Proposed method includes determination of charge-discharge characteristics of onboard power supply sources, transformation of energy of external power sources, maintenance of required state of charge of onboard power sources due to energy of external power sources and consumption of energy power requirements exceed transformed energy of external power sources. Intervals of flight time required for maintenance of probable and determined level of state of charge of onboard power sources are also determined. At probable level when power requirements exceed transformed energy of external power sources, amount of energy in onboard power source required for performing the program is determined. Then, intervals of time of charge and self-discharge of onboard power sources are determined; besides that, shift of beginning of charge of onboard power sources ensuring required amount of energy is determined. Required state of charge of onboard power sources is maintained at said intervals of their charge-discharge at control of power consumption in accordance with flight program. When energy transformed from external power sources exceed consumed power, onboard power source is charged and state of charge is maintained at charge-discharge intervals. Time intervals for maintenance of determined state of charge of onboard power sources when consumed power exceeds transformed power of external sources is predicted. Before beginning of these intervals, determined charge of onboard power sources is performed at said charge-discharge intervals. Then, power requirements of onboard power sources are regulated in accordance with flight program ensuring excess of discharge energy of onboard power sources by consumed energy. Upon completion of predicted interval, onboard power source is charged for subsequent maintenance of probable state of charge of onboard power sources. When necessary shift is made for above-indicated determined charge of onboard power sources after which probable and determined charge are alternated. |
Another patent 2550831.