Solar battery for small-size spacecrafts and method of its manufacturing |
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IPC classes for russian patent Solar battery for small-size spacecrafts and method of its manufacturing (RU 2525633):
  
 Polymer photovoltaic module and method for production thereof / 2519937
Polyaniline is doped with a heteropolyanionic complex of the 2-18 series, having chemical formula [P2W18O62]6-. A doped polyaniline film 1 is deposited on a thin transparent conducting layer which may consist of indium (III) oxide or tin (IV) oxide 2, which in turn is sputtered onto a material 3, having high transmission capacity for electromagnetic waves in the range from 3·10-2 to 4·10-6 cm. Said material with the sputtered conducting layer and the polyaniline film forms one of the electrodes of the photovoltaic module, and a second counter electrode, which also serves as back wall of the article, can be made of conducting material 4, on the outer side of which are attached thermogenerators 5 with air or water radiators for removing heat 6, connected to each other by series-parallel electrical circuits 7, and the electrodes held with each other by side walls, which can be made of any non-aggressive dielectric material 8, and an aqueous electrolyte is poured between the electrodes, where pH of the electrolyte 9 may vary from 5 to 3; current terminals are respectively attached to the conducting material with the polymer film and to the conducting back wall of the article, and to output terminals of the thermogenerators 10 to form two independent electrical circuits. The invention also relates to a method of producing said module.
 Grate of photogalvanic cells with mechanical detachment of cells relative to their support / 2518021
Use: to implement solar generators panels to ensure electricity supply of spacecrafts, in particular satellites. Essence of invention consists in the fact that each photogalvanic element of the grate is mounted on the substrate using the soft and self-adhesive and easily detachable fastening device, at that the rear side of each cell and the front side of the substrate are coated with a layer that improves their properties of heat radiation.
 Silicon multi-junction photoelectric converter with inclined structure and method for production thereof / 2513658
Present invention relates to silicon multi-junction photoelectric converters of solar cell panels. The structure of an "inclined" silicon monocrystalline multi-junction photoelectric converter according to the invention includes diode cells with n+-p--p+ (p+-n--n+) junctions which are parallel to a horizontal light-receiving surface; the diode cells include n+(p+) and p+(n+) regions of n+-p--p+(p+-n--n+) junctions through which they are connected into a single structure by metal cathodes and anodes placed on the surface of n+(p+) and p+(n+) regions to form corresponding ohmic contacts - connections, wherein the n+(p+) and p+(n+) regions and corresponding cathodes and anodes are placed at an angle in the range of 30-60 degrees to the light-receiving surface; the metal cathodes and anodes are placed on their surface partially, and partially lie on the surface of an optically transparent dielectric which is placed on the surface of n+(p+) and p+(n+) regions, wherein they form an optical reflector with the metal electrodes and the optically transparent dielectric. Also disclosed is a method of making the described structure of an "inclined" silicon monocrystalline multi-junction photoelectric converter.
 Combined production of heat and electric energy for residential and industrial buildings with application of solar energy / 2513649
In accordance with the invention claimed solar-powered generator (100) contains thermoelectric elements adjoining solar elements and located below solar elements. Concentrated flow of solar energy is provided. Heat sink (104), with changeable temperature and efficiency, contacts with cold soldered seam (108) of thermoelectric device (103). Thermal resistance is calculated with respect to energy flow, which results in creation in thermoelectrical device (103) of temperature gradient equal to several hundreds of Kelvin degrees. Solar element preferably contains semiconductor with large width of prohibited energy zone. Generator (100) preserves relatively suitable efficiency (efficiency factor) in some range of cold seam (108) temperature. System of hot water can serve as heat sink (104). High values of efficiency factor are obtained due to application of nanocomposite thermoelectrical materials. One-piece construction of solar element and thermoelectrical elements provides additional advantages.
 Multipurpose solar power plant / 2505887
Multipurpose solar power plant (hereinafter referred to as MSPP) refers to renewable power sources, and namely to use of solar radiation to generate electric power, provide hot water supply and natural illumination of rooms of different applications, which contains the following: an optically active transparent dome representing a rectangular biconvex lens, a photovoltaic panel, a solar collector, round flat horizontal dampers of hollow light guides, hollow light guide tubes, a heat-receiving copper plate of the solar collector, a solar light dissipator, micromotors of round flat horizontal dampers of hollow light guide tubes, round light-emitting-diode lamps, storage batteries, light and temperature sensors, an electronic control unit, a control panel, a storage tank, a heat exchanger, a pump, a check valve, six-sided copper pipelines, an inverter and a support with support racks to support MSPP structure.
 Solar module with concentrator (versions) and method of its manufacturing / 2503895
In a solar module with a concentrator comprising a transparent focusing prism with an angle of complete inner reflection
 Making solar cell modules / 2501120
Disclosed is use of a) polyalkyl(meth)acrylate and b) a compound of formula (I), wherein residues R1 and R2 independently denote an alkyl or cycloalkyl with 1-20 carbon atoms, to make solar cell modules, primarily for making light concentrators of solar cell modules.
 Back sheet for solar cell module and solar cell module / 2498458
Back sheet for a solar cell module has a substrate sheet and a cured layer of a coating film made of coating material, formed on one side or each side of the substrate sheet, wherein said coating material contains a fluoropolymer (A), having repeating units based on fluoro-olefin (a), repeating units based on a monomer (b) which contains a cross-linking group, and repeating units based on a monomer (c) which contains alkyl groups, where the C2-20 linear or branched alkyl group does not have a quaternary carbon atom, and unsaturated polymerisable groups are bonded to each other through an ether bond or an ester bond. Also disclosed is a solar cell module using said back sheet and versions of a method of making the back sheet for the solar cell module.
 Flexible photoelectric module / 2495513
Flexible photoelectric module comprises the following serially arranged components: a lower bearing film, a lower reinforcing net, a lower fixing film, electrically connected solar elements from single-crystal silicon, an upper fixing film, an upper reinforcing net and an upper bearing film. The bearing and fixing films are made of a material, which is transparent for sun light, and reinforcing nets are made of polymer threads, which are transparent for sun light and are impregnated with a substance or containing such substance with low coefficient of light absorption and scattering. Reinforcing nets are annealed nets from a thermosetting polymer.
 Semiconductor photoelectric generator (versions) / 2494496
Semiconductor photoelectric generator with double-sided working surface is made as a matrix from switched microphoto cells with n+-p-p+(p+n-n+) diode structures, in which one or two linear dimensions of the microphoto cell are comparable with diffusion length of minor current carriers in the base area, and planes of diode structures are inclined at the angle φ, 30°<φ<150° to the working surface of the generator, along the entire area of the working surface at two sides of the generator there is a passivating film with thickness of 10-60 nm, arranged on the basis of one or two oxides of the following metals: tantalum, zinc, aluminium, molybdenum and tungsten, and above the passivating film there is a layer of a clearing coating. In the other version along the entire area of the working surface of the generator at two sides of the generator there are passivating and clearing films, made on the basis of one or two oxides of the following metals: tantalum, zinc, aluminium, molybdenum and tungsten, and also silicon nitride or carbide.
 Method of control over spacecraft solar battery orientation with limitation of solar battery turn angle / 2509694
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.
 Method of control over spacecraft solar battery orientation with control over spinning direction and continuous change of data on solar battery angular position / 2509693
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.
 Method of control over spacecraft solar battery with protection against short-term faults of data on solar battery angular battery / 2509692
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.
 Solar battery strut / 2499751
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.
 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.
 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.
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FIELD: electricity. SUBSTANCE: result is achieved by increasing strength of connection of shunting diodes and solar elements, increased repeatability of the process of manufacturing of the solar battery of spacecrafts due to optimisation of the technology of manufacturing of shunting diodes and solar elements of the solar battery, and also switching buses that connect the solar elements and shunting diodes, which are made as multi-layer. The solar battery for small-size spacecrafts comprises the following: panels with modules with solar elements (SE) adhered to them, a shunting diode; switching buses that connect the face and reverse sides of the shunting diodes with solar elements, at the same time the shunting diode is installed in the cut in the corner of the solar element, at the same time switching buses are made as multi-layer, made of molybdenum foil, at two sides of which there are serial layers of vanadium or titanium, a layer of nickel and a layer of silver, accordingly. EFFECT: increased resistance of solar batteries to thermal shocks, to impact of mechanical and thermomechanical loads, increased manufacturability of design, extended active life of a solar battery of spacecrafts, increased functional capabilities due to expansion of temperature range of functioning and optimisation of design of a solar battery, simplification of a switching system. 7 cl, 4 dwg, 3 tbl
The technical field The invention relates to electrical engineering, in particular to a device for generating electrical energy by converting light radiation into electrical energy, and can be used in the creation and production of small spacecraft with solar batteries (SB). The level of technology To SB must meet the following requirements: maximum energy efficiency with minimum weight, maintaining electrical and mechanical characteristics during storage, transportation on the Ground and the output into a geosynchronous transfer orbit, long lifetime (CAC) in orbit with minimal degradation, which is expressed in loss of power. Modern SAT CAC reaches 15 years of age and requirements on its extension up to 20 years. The main causes of degradation in orbit are disrupting the structure of the active elements, namely photoconverters (AF) and diodes under the action of radiation, as well as violations resulting from the impacts of a change in temperature of thermal cycles. On different orbits varies the range of temperature and frequency of thermal cycles. For operating conditions in a geostationary orbit, the upper temperature +100°C, the lower - 170°C, number of cycles - 2000. On low orbits the range of changes in temperature, m is nice, the upper value +100°C, lower - 100°C, but the number of cycles during the active lifetime in orbit is several tens of thousands. In the prior art it is known (see N. S. Rauschenbach. The principles and technology of photovoltaic energy conversion. New York, 1980)that SAT consists of separate generators, including chain of solar cells (SCS)in the generators of the counter-parallel with the solar cells installed bypass diodes. In addition to the bypass diodes to ensure reliable operation of the SAT is used diode protection that is provided by a blocking diode. In recent years, to replace silicon came more efficient solar cells, which includes several stages of heterojunctions based compounds ASV grown on a germanium substrate (see P. R. Sharps, M. A. Stan, D. J. Aiken, W. Clevenger, J. S. Hill and N. S. Fatemi. High efficiebcy, multi-junction cells with monolithic bypass diodes, NASA/CP.2005-213431. Page 108-115). Each solar cell is protected by a diode, located on the SE in the same plane, and the diode has the same thickness as and SE. Usually solar cell made according to the angles of the slices in which the diode is triangular in shape (see U.S. patents for inventions US 6353176, US 6034322 and application of the U.S. for the invention of US 2008/0000523). The prior art solar cell spacecraft, located on the carbon honeycomb. Supporting part cell panel which consists of two layers of carbon fiber, between which is located a cellular filler of aluminum foil. On the carbon surface, designed for installation of SCS, glued insulating film. Energy-generating part of the solar panels (modules) consists of solar cells in series or series-parallel connected to each other through switching elements to thermo-mechanical expansion joints. On the front surface of each solar cell is glued to a glass plate (see GLOBASTAR. Solar Generator Desigh And Layout For Low Earth Orbit Application in Consideration Of Commercial Aspects And Quanlity Production. D-81663 Munich Germany). The disadvantages of the known solar spacecraft are low manufacturability of the design, a small temperature range of operation due to the low strength of soldered and welded joints bypass diodes and solar. High probability of damage to the inter-element switching, protruding over the front surface of SAT, in the process of its production and maintenance, as well as the technological complexity of manufacturing inter-element switching, caused by the necessity of placing thermo-mechanical expansion joints in narrow inter-element gaps, leads to low SB resistance to the effects of thermal and mechanical loads. The closest in technical essence and the achieved effect is KTU technical solution (prototype) of a solar battery of the spacecraft, contains panels with glued on them modules, consisting of a series or series-parallel connected with the switching Shin SE, where the switching bus is equipped with thermo-mechanical expansion joints, and to the front surface of each solar cell bonded protective glass plate, which is provided in addition glued to flat or curved surfaces of the frame having a specified shape and size of the elastic elements, where the internal volume of the elastic elements is filled with a sealant with the formation of a convex meniscus, and SE is pressed against an elastic elements and fixed motionless, and switching bus to thermo-mechanical expansion joints and shunt diodes are welded or soldered to the back contact solar cell in the free zones from the sealant, and thermo-mechanical expansion joints are located between the back of the AOC and the bearing surface of the frame in areas free from sealant (see Russian Federation patent for the invention EN 2250536). The disadvantages of the known solar spacecraft are low manufacturability of the design, a small temperature range of operation due to the low strength of soldered and welded joints bypass diodes and solar, bad SB resistance to mechanical and thermo-mechanical loads. Molybdenum Chi is a, thickness of 50 μm and having a special multilayer coating is very hard. When joining a commuting tire welding deteriorate electrical characteristics of the bypass diodes, and in some cases due to the hard tyre point of welding is pulled out together with the silicon, resulting in low yield of the crystals after testing thermal Cycling. At elevated temperatures, there is degradation of the solar cell after soldering and welding, which leads to peeling of the contact from the AOC and, as a consequence, the output of the operational state of the cells SAT. The prior art method of making SB spacecraft with bypass diode, which includes the production of SE-based photovoltaic semiconductor substrate, forming a bypass diodes on the front side of the solar cell, the connection of shunt diodes and SCS SB spacecraft, the connection through the switching Shin SE (see U.S. patent for the invention US6635507). The disadvantages of this process is the low reproducibility of the manufacturing process due to the high probability of damage (loss of adhesion) of the metallization on the working and non-working sides. In addition, when attaching a commuting tire welding possible circuit switching bus layer structure, and the point of welding is pulled out together with the structure of the substrate, which leads, as the trail is a journey, low yield of crystals after testing thermal Cycling. The closest in technical essence and the achieved effect technical solution (prototype) is a method of manufacturing SB spacecraft with integrated bypass diode, which includes the production of SE-based photovoltaic semiconductor substrate with grooves to accommodate discrete shunt diodes, discrete manufacturing bypass diodes based on semiconductor substrate, mounting a discrete bypass diodes in the recesses, the contact solar cells with bypass diodes with switching buses (see U.S. patent for the invention of US 5616185). The disadvantages of this method of manufacture is the low reproducibility of the manufacturing process due to the high probability of damage (loss of adhesion) metallization during the formation of metallization outside parties. In addition, when cutting on the crystal on silicon single crystal substrates cracks, and the accession switching buses welding point of the welding escapes together with silicon, which leads consequently to a low yield of crystals after testing temperature Cycling (thermal shock). Disclosure of inventions The technical result of the claimed invention is the I: - increased durability SAT to thermal shocks, mechanical and thermo-mechanical loads, increase of technological design, increasing the lifetime SAT spacecraft, increase functionality by expanding the temperature range of operation and optimization of design SAT, - simplification of the switching system, which is achieved by increasing the strength of the connection of the bypass diode and the solar cell - improving the reproducibility of the manufacturing process SAT spacecraft for optimizing manufacturing technology shunt diodes and solar SA, as well as commuting tires, connecting solar cells and bypass diodes, which are made of multilayer. The technical result of the claimed invention is achieved by the fact that the solar battery of small spacecraft contains: panels with glued on them modules with solar cells (SCS), - shunt diode; - commuting tires, welded to the front and back sides of the bypass diodes and connecting the front and back of the bypass diode with the AOC, with the shunt diode is mounted in a cutout in the corner of the AOC while commuting tires are made of a multilayer consisting of molybdenum foil on both sides of the sequentially applied layer VA is pushing or titanium, a layer of Nickel and a layer of silver, respectively. In a preferred embodiment, the thickness of the molybdenum foil is 8-12 μm, the total thickness of the layers of vanadium or titanium and Nickel is 0.1-0.3 μm, the thickness of the silver layer is 2.7-6 ám. A method of manufacturing solar panels for small spacecraft includes: - manufacturer of solar cells (SCS) - based photovoltaic semiconductor substrate with a cutout in the corner under shunt diodes - production of bypass diodes on the basis of the photoelectric semiconductor substrate, - manufacturer of switching tires - welding of the switching tires to the front and back sides of the shunt diodes - installation of bypass diodes in the cutout in the corner of SE, connection SE shunt diode by means of a commutating tires while commuting tires are manufactured from molybdenum multilayer foil, with two sides which consistently put a layer of vanadium or titanium, a layer of Nickel and a layer of silver, respectively. In a preferred embodiment, the layer of vanadium or titanium, a layer of Nickel and a layer of silver is applied in series with the two sides prepared on molybdenum foil by vacuum magnetron sputtering at a temperature of molybdenum foil 110-130°C With pre ion bombardero the coy, and the molybdenum foil formed with layers of vanadium or titanium, Nickel and silver annealed in vacuum at a temperature of 300-350°C. Brief description of drawings Signs and essence of the claimed invention are explained in the following detailed description, illustrated by the drawings, there is shown following. Figure 1 shows the solar cell installed on the side with the switching tires shunt diode. Figure 2 is a schematic representation of the layered structure of the switching tires. Figure 3 presents the algorithm of the method of manufacturing SA KA. Figure 4 presents calculated from experimentally measured deformation values of the internal stress in the metal layers of the switching tires, formed at various temperatures of molybdenum foil. Figure 4 on the charts in brackets optimal operating temperature range molybdenum foil during the deposition. Figure 1 indicated the following: 1 - shunt diode; 2 - commuting bus, which connects the front side of the shunt diode(1) SE (4); 3 - commuting bus connecting the opposite side of the shunt diode(1) SE (4); 4 - solar cell (solar cell); Figure 2 indicated the following: 5 - prepared molybdenum foil; 6 - layer vanadium or titanium; 7 - layer of Nickel; 8 - layer silver. The implementation of the representation and an example implementation of the invention The claimed method was used for the implementation of group technology manufacturing solar panels of spacecraft and consists of the following sequence of technological operations (see figure 3): produce solar cells based photovoltaic semiconductor substrate, the manufacturing bypass diodes on the basis of the photoelectric semiconductor substrate, making commuting tire, including the preparation of the molybdenum foil and the metallization prepared molybdenum foil by vacuum magnetron sputtering from two sides by layers of vanadium, Nickel and silver at a temperature of molybdenum foil 110-130°C with pre ion bombardment, and then produce the annealing molybdenum foil formed with layers of vanadium or titanium, Nickel and silver vacuum at a temperature of 300-350°C, carry out welding of the switching tires to shunt the diodes have bypass diodes for temperature Cycling and thermal shock, connect the solar cells with bypass diodes with switching tires and perform output control of the solar panels of the spacecraft. The thickness of the molybdenum foil was chosen based on the highest breakout force is welded commuting tires to the front and back sides of the shunt diode after about what edenia test thermal shock. Breakout force is welded to the switching bus from the bypass diode was defined as follows: preparing a molybdenum foil in several stages, after which were thinning the molybdenum foil to the following thicknesses: 6±0.1 ám and 7.5±0.1 ám, 10±0.1 ám, 13±0.1 ám. Then prepared on molybdenum foil by vacuum magnetron sputtering has dealt with the two sides of the layers of vanadium, Nickel and silver at a temperature of molybdenum foil 110-130°C with pre ion bombardment. Then molybdenum foil formed with layers of vanadium or titanium, Nickel and silver were annealed in vacuum at a temperature of 300-350°C and produced the felling of the molybdenum foil commuting tires. After which he made reference welding switching tires to the front and back sides of the bypass diodes and control breakout force commuting tires from bypass diodes (see table 1).
Then tested for thermal shock welded commuting tires to shunt diodes, which consists in holding 450 cycles of a thermal shock temperature -180°C (a pair of liquid nitrogen) to 120°C on specialized equipment. After that was done the measurement of electrical parameters of bypass diodes, which showed a slight increase in forward voltage on the background of unchanged values of leakage current and reverse voltage. Then made the control breakout force commuting tires from bypass diodes (see table 2). 100-125
The test results revealed an increase in the effort of separation in all cases, the thickness of the switching tires from bypass diodes with little change in electrical characteristics of the bypass diodes. From table 2 it is found that the optimal thickness of the molybdenum foil is 10±0.1 ám, as it provides maximum pullout force bus from the bypass diode. The temperature of the molybdenum foil with the technological operation of the spraying of metals were chosen on the basis of the minimum stress in the resulting structure (see figure 4). Internal stress was determined as follows: formed odnosezonnye microbulk by magnetron sputtering of metal films V-Ni-Ag prepared on molybdenum foil by photolithography and chemical etching of metals. The samples dnacontaining microblock was investigated using an optical microscope Axio Imager, Carl Zeiss with the increase in h. Measurements of the dimensions of the beam structure and the direction of deformation. Shape deformation was determined by the deviation of the microbe is Lok at different points of its length from the surface. Then using a mathematical processing by the formula stony calculated voltage beams. The curvature of the beams were found by measuring the deviation of the shank odnokomnatuyu microbulk. These modes were chosen for reasons of reproducibility of the process, which is provided, if accession commuting tyres welding point of the welding does not burst (see table 3).
On the proposed design and methods of manufacture produced SB for small SPACECRAFT, including a frameless bypass diodes triangular reverse the voltage of 100 V and direct current of 2 a and cascade solar cells based compounds And 3In5. Prior to the use of the claimed technical solution used a silver patch of the tires, which had to be welded to the clamp diodes and solar. Test diodes showed low resistance to thermal shocks (was the destruction of the structure after 10-15 thermal shocks from -180°C to +100°C), and the percentage of yield diodes electrical characteristics at the stage of thermal Cycling was less than 70% of suitable diodes after Assembly, and the remaining 30% was the destruction of structures in the zone of welding (interlayer destruction on the basic materials when exposed to high and low temperatures) when testing the strength of the welded joint. The breakout metallization from the crystal was 50-100 g/mm2and after using this technical solution amounted to more than 150 g/mm2, resulting in the percentage of yield of diodes at the stage of thermal Cycling increased to 85%. 1. Solar panel for small spacecraft contains: 2. Solar battery according to claim 1, characterized in that the thickness of the molybdenum foil is 8-12 microns. 3. Solar battery according to claim 2, characterized in that the total thickness of the layers of vanadium or titanium and Nickel is 0.1-0.3 microns. 4. Solar battery according to claim 3, characterized in that the thickness of the silver layer is 2.7-6 ám. 5. A method of manufacturing solar panels for small spacecraft, including: 6. the procedure according to claim 5, characterized in that the layer of vanadium or titanium, a layer of Nickel and a layer of silver is applied in series with the two sides prepared on molybdenum foil by vacuum magnetron sputtering at a temperature of molybdenum foil 110-130°C with pre ion bombardment. 7. The method according to claim 6, wherein the molybdenum foil layers formed of vanadium or titanium, Nickel and silver annealed in vacuum at a temperature of 300-350°C.
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