Solar battery strut |
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IPC classes for russian patent Solar battery strut (RU 2499751):
           
 Space solar electric station and independent photo emitting panel / 2492124
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.
 Solar cell battery / 2485026
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.
 Bench for opening panels of solar battery / 2483991
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).
 Solar battery drive system / 2466069
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.
 Method of spacecraft solar battery position control during partial failures of aspect sensor / 2465180
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.
 Method of spacecraft solar battery orientation by electric current / 2465179
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.
 Device for electromechanical linkage between fixed structural parts of spacecraft / 2462400
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.
 Solar cell panel for spacecraft / 2460676
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.
 Method for space vehicle with fixed panels of solar batteries orientation control during experiments on orbits with maximum eclipse period / 2457158
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.
 Unit offices and opening the valves of solar batteries of spacecraft / 2441817
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.
 Method of control of power of spacecraft power plant and device for realization of this method / 2249546
Proposed method includes stabilization and change of power of power plant through regulation of consumption of engine working medium. When power of solar battery drops to level of maximum permissible power consumed by engine, consumption of working medium is changed in such way that power of solar battery might change in saw-tooth pattern and vertices of saw might be in contact with line of maximum probable power of solar battery. Device proposed for realization of this method includes matching voltage converter whose outputs are connected with engine electrodes and inputs are connected with solar battery busbars, current and voltage sensors showing solar battery voltage and power sensor connected with current and voltage sensors. Comparator connected with power sensor is also connected with controllable power setter and initial power setter. Outputs of controllable power setter are connected with comparator and comparison circuit whose input is connected with power sensor output. Output of comparison circuit is connected with amplifier-regulator of consumption of working medium.
 Solar battery (versions) / 2258640
Proposed solar battery includes panels foldable by "bellows" pattern and frame with drive mechanism. Panels are interconnected together and are connected with spacecraft through frame by means of drive springs and cable run with pulleys. Provision is made for articulated rods of adjustable length and locking units for locking the solar battery in folded and open positions and several contact components (pins, seats, pushers) for interaction of solar battery panels. Locking units are made in form of stops with elliptical holes and spring-loaded retainers. According to first version, solar battery includes fixed pulleys on spacecraft, intermediate pulleys on first panel of solar battery, brace movable relative to spacecraft, strut movably connected with frame and with brace (through spring) and stop engageable with brace. According to other version, drive mechanism is provided with engine and pulley connected with intermediate pulley by means of cable run. Engine and pulley are secured on spacecraft by means of bracket with elliptical locking holes. Movable unit of engine is fastened with frame and frame is provided with spring-loaded pins locked in locking holes of bracket in opening the solar battery.
 Method of forming thrust in solar radiant flux and device for realization of this method / 2268206
Proposed method includes forming light-sensitive surface of solar sail and orientation of this surface in solar radiant flux. This surface is formed as cloud of finely-dispersed particles charged by solar photoelectrolizing. Stable shape close to sloping surface is imparted to cloud by means of electrostatic system of spacecraft. This system has at least one central and one concentric charge carriers of opposite signs. Control of shape and sizes of cloud may be performed by screening central charge or moving it relative to circular charge.
 Solar sailing craft / 2269460
Proposed craft has hull, main and additional circular reflecting surfaces, units for forming such surfaces provided with twisting devices and control units for orientation of these surfaces. Orientation control units are made on base of gimbal mounts brought-out beyond craft hull. Each twisting device is made in form of hoop mounted on outer frame of gimbal mount for free rotation; it is engageable with electric motor. Units for forming reflecting surfaces are made in form of pneumatic systems with concentric pneumatic chambers and radial struts. Said struts are provided with flexible tubes with valves mounted at equal distances. Valves have holes. Built on said tubes are pneumatic cells in form of torus or spheres. Each pneumatic system is mounted on respective hoop and is communicated with compressed gas source through concentric hermetic groove found in hoop and in outer frame of gimbal mount.
 Foldable and unfoldable complex of components mounted on board of spacecraft / 2271318
Proposed complex contains components (1.1-1.n) rigidly connected with side (3) of soft inflatable mat (4). In transportation position of components, mat (4) is in deflated state and is folded in such way that components of complex are located on both sides of fold (5.1-5.n-1) of mat in pairs.
 Spacecraft / 2271965
Proposed spacecraft is equipped with solar sail, central fixed module and movable module which is coaxial relative to first module and is provided with bio-energy complex. Laid spirally on surface of movable module are growth tubes with plant conveyers which ensure turn of movable module around central axis. Connected with modules are generator and electric power accumulator. Fixed module is provided with cylindrical separable ice melting modules. Each module is provided with parachute for descent on planet, its own bio-energy complex and ice melting chamber for forming shaft in ice cover of planet. Ring of reactors located around central axis of module are combined with toruses. On side of central axis reactors are coated with warmth-keeping jacket and are provided with heaters and units for filling the reactors with water in lower part and with oil in upper part. These units ensure operation of hydraulic generator generating vapor for melting ice and supplying distilled water to bio-energy complex. Modules are provided with envelope pressurizing units, deploying their parachutes and supplying sea water from shafts to envelope surfaces for forming ice domes. When domes are combined, stations may be formed for research of planet followed by its populating. Modules are equipped with descent bathyspheres for research of under-ice ocean and robots for performing jobs on planet surface. Spacecraft may include manned separable raiders and bathyscaphs for research of ocean depth. Both of them may be provided with their own bio-energy complexes.
 Spacecraft with power supply units / 2271968
Proposed spacecraft has form of right-angle prism with cross-section in form of equilateral tetragon (rhomb). Mounted on side faces of prism are solar battery panels. Spacecraft is provided with passive or combined system of gravitational stabilization in orbit. Acute angle of tetragon ranges from 50 to 90° to ensure required power supply for spacecraft equipment. Main central axes of symmetry of spacecraft in transversal plane are parallel to tetragon diagonal. Lesser axis is parallel to larger diagonal, thus enhancing stable gravitational orientation of spacecraft by larger diagonal perpendicularly to orbit axis.
 Method of control of cluster of satellites in geostationary orbit (versions) / 2284950
Proposed method consists in measurement of parameters of satellite orbits, determination of orbital elements, comparison of them with required ones and performing of correcting maneuvers with the aid of thrusters. Satellite inclination vectors are brought to circular areas of their permissible change which are spaced apart so that angle between line connecting the end of vector with center of its circular area and direction to Sun should exceed right ascension of Sun by 180°. According to first version, vectors of satellite eccentricity are shifted to similar circular areas so that similar line lags behind direction to the Sun by half angular displacement of vector over circumference of its natural drift within circular area. Then, distances between satellites are changed within required limits compensating for quasi-secular increment of inclination vector and correcting eccentricity vector so that at passing the center of interval between point of circumference entry of its natural drift to its circular area and point of exit from this area, line connecting the center of this circumference and center of circular area coincide with direction to the Sun. In case circular area of permissible change of each eccentricity vector is close to circumference of its natural drift (second version), said line for this area is matched with direction to the Sun and no correction is made in this case.
 Concentrating solar-electric generator / 2285979
Proposed solar generator module has at least one cellular-structure panel 1 incorporating front face sheet, rear face sheet, and cellular lattice in-between. Front sheet mounts alternating rows of solar cells 2 and wedge-shaped reflectors 3. The latter may be of developable type, for instance made of thin film stretched on stiff frame which do not cover solar cells 2 in folded condition. One of generator-module design alternates may have additional cellular-structure lattice attached to rear face sheet. At least one of face sheets is made of polymer incorporating high-heat-conductivity threads positioned in average perpendicular to longitudinal axis of rows of solar cells 2. Module may incorporate at least two hinged cellular panels folded along hinge whose reflectors 3, for instance non-developable ones, are alternating in folded condition without contacting each other. Panel mechanical design affords maintenance of uniform sun radiation distribution among all cells of generator module at small deviations from sun rays. Reflectors may be covered with aluminum layer or better silver one applied by vacuum evaporation and incorporating additional shield.
 Method of control of spacecraft power supply system / 2291819
Proposed method includes conversion of light energy into electrical energy on board spacecraft, accumulation of electric energy by conversion into other kinds of energy and check of onboard power requirements. In addition to conversion of light energy, other kinds of energy received from outside sources are transformed into energy consumed on board aircraft. Provision is made for prediction of time intervals when consumption of electric energy exceeds amount of electrical energy converted from light energy. Kinds of consumed energy obtained from conversion of electric energy are also determined. Before beginning of passage of intervals, respective transformable kinds of energy from outside sources are accumulated. Beginning of accumulation of these kinds of energy is determined depending on consumed amount of energy and rate of accumulation of energy with change in spacecraft parameters caused by action of these kinds of energy taken into account. At predicted interval, first of all accumulated transformable kinds of energy are consumed and when necessary, energy accumulated on board spacecraft is consumed after conversion into electrical energy.
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FIELD: aircraft engineering. SUBSTANCE: 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. EFFECT: higher reliability, simplified installation of solar battery at craft body. 13 dwg
The invention relates to space technology and can be used when designing constructions of space vehicles (SV), mostly antennas and solar panels. Known solar battery containing directly sash, frame and brace, movably mounted relative to the spacecraft, rack movably connected to the frame and strut with coil springs, interacting with the strut, and a stopper made with the possibility of interaction with the knee after moving the axis of rotation of the rack relative to the frame behind the plane formed by the axes of rotation of the mentioned rack relative to the strut and the strut relative to the spacecraft (patent RU №2258640, MPK7 B64G 1/44 - equivalent). The disadvantage of this design is the low reliability due to the presence of wire rods and spring tension placed on the outer surface of the strut, as well as opportunities folding of the strut in the opposite direction due to the absence of the release end position and appearance of shock loads at the time of disclosure of the strut into its final position. Known the knee of the solar battery 1, containing two-link mechanism, the link 2 which are mounted on the frame 3 of the solar battery 1 and link 4 is installed on the housing 5 of the spacecraft, and on the axis connecting the two link, set the go torsional spring 6 devices cocking 7 and 8, spring-loaded rod 9, which locking element in the end position of the knee and the solar battery 1, for which disclosure is drained from the housing 5 of the spacecraft, turning around the axis 10. When the abstraction of the solar battery 1 from the housing 5 spacecraft spring-loaded rod 9 by means of a spring 11, overcoming the sliding friction, moves along the cylindrical surface 12 of the element 2 to hit into the hole 13 of larger diameter on the surface, under the action of the torsion springs 6, cocked devices 7 and 8, the element 2 comes to the end position with the stroke adjusting bolt 14 in the side surface 15 of the link 4, and a spring-loaded rod 9, falling into the hole 13 on the cylindrical surface 12 of the element 2 forms a gap 16 (working drawings F-SB, F-SB, TsSKB, gcabashe, 1981 - prototype 1, 2, 3, 4). Figure 1 - General view of the installation of solar panels on the hull of the SPACECRAFT. Figure 2 - General view of the design in the original position. Figure 3 - General view of the structure in its final position. 4 is a View As in figure 3. The disadvantages of this design of the knee are the presence of sliding friction, the presence of shock loads on the structure of the solar battery, and the impossibility without shame placement of the locking element in the end position makes it impossible to install the strut on exactly the problem is, Mr. angle disclosure, thereby reducing the reliability of the structure during the abstraction of the solar battery from the product, and the installation of the strut in position requires additional labor, which complicates the process of installing solar panels on the hull of a spacecraft. The present invention is to remedy these disadvantages, namely improving the reliability of the knee by eliminating sliding friction, the exception shock loads and accurate recording end position of parts of the knee when the abstraction of the solar battery from the device at a certain angle, as well as simplify the installation of solar panels on the hull of a spacecraft. The problem is solved in that in the proposed design, containing two-tier mechanism, and one link is installed on the frame of the solar battery, and the other on the body of the spacecraft, for a total of two links of the axis which has a torsional spring with your cock, and perpendicular to this axis on one of the links is spring-loaded rod to lock in its final position at the end of the spring rod can be rotated installed the rocker on both ends of which is rigidly fixed to the bearings, interacting with tapered grooves copiers rigidly mounted nepretvorennom podpruzhinennom the stock link, and in parts of the two-tier mechanism made hole device for fixing the initial position of the links attached by threaded connection. The claimed design is illustrated by drawings: 5 is a General view of the proposed design in the original position. 6 is a View In figure 5. 7 is a General view of the proposed design in end position under a certain (predetermined) angle. Fig - remote element D figure 6. Fig.9 - Type E figure 7. Figure 10 - cross Section G-g of figure 6 (initial position). 11 - Section C-C on Fig (original position). Fig - Section K-K figure 9 (end position). Fig - Section And figure 9 (end position). The design of the strut consists of a two-tier mechanism consisting of links 17 and 18 (figure 5), and the link 17 can be rotated in the direction B relative to the link 18, with a common axis and mounted on the torsion springs 19 and 20, the ends of which are devices cock 21 (6), and spring rod 22 by a spring 23 (figure 10), at the end of which can be rotated is the rocker 24 is rigidly mounted on the bearings 25 (Fig), which in turn communicate with the master plates 26 and 27 are rigidly mounted the link 18 (11). Each of the copiers 26 and 27 has a radial plot, lane is walking in the tapered groove 28. The links 17 and 18 in the initial position under the action points from the torsion spring 19 and 20 are held by the locking device 29 (figure 10), which is situated in the smooth bore 30 of the link 18, is screwed into the threaded hole 31 of the link 17, which provides the necessary angle between the links 17 and 18 of the strut and greatly simplifies the installation of solar panels on the product. The operation of the lock is as follows. Brace screwed with the locking device 29 is installed between the solar panel and the body of the spacecraft. The locking device 29 come unscrewed, and the links 17 and 18 occupy its original position. When the abstraction of the solar battery from the body of the spacecraft in the knee is the rotation of the links 17 and 18, which under the action of torsion springs 19 and 20 help solar battery to take the final (working) position. During rotation of the links 17 and 18, the bearings 25, moving on the radial surface of copiers 26 and 27 in the direction W, by means of a spring-loaded rod 22 by means of a spring 23 (Fig) fall into the corresponding tapered grooves 28 of these copiers in the direction L (Fig) and fix the link 17 link 18 at a given angle in our case 179° (Fig.7). Thus, the proposed solution will allow to increase the reliability of knee solar panels by eliminating t is possible slip and gap, which in turn eliminates shock loads, and to simplify the process of installing solar panels on the hull of a spacecraft. Brace solar battery containing two-tier mechanism, and one link is installed on the frame of the solar battery and the other on the body of the spacecraft, for a total of two links of the axis which has a torsional spring with your cock, and perpendicular to this axis on one of the links is spring-loaded rod for locking in the end position, characterized in that at the end of the spring rod can be rotated installed the rocker on both ends of which is rigidly fixed to the bearings, interacting with tapered grooves copiers, rigidly mounted on the opposite podpruzhinennom the stock link, and the links of the two-tier mechanism made hole the device for fixing the initial position of the links attached by threaded connection.
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