What solar panels are used in space. Solar battery for small spacecraft and method for its manufacture. Spacecraft of the Luna mission

At present, NPP Kvant is working on three main areas of development of space photovoltaics and its elemental base, namely:

Creation of solar cells based on single-crystal silicon

Silicon solar batteries created at NPP Kvant correspond to the world level, which was confirmed during the fulfillment of a number of foreign orders for their manufacture in the interests of India, France, Holland, the Czech Republic, Israel, and China. These batteries have:

• the highest initial specific energy characteristic ~ 200W/m 2 ;
• the least degradation over the period of active existence;
• two-way sensitivity, which is used on low-flying spacecraft and allows you to increase the output power of solar panels by 10-15% due to the transformation of the Earth's albedo (in particular, solar batteries for the Zarya, Zvezda spacecraft, the Russian sector of the ISS, SB for the spacecraft " Monitor-E").

Creation of solar batteries based on multistage photoelectric converters using complex semiconductor materials on foreign substrates.

With the help of solar cells based on cascaded complex heterojunction structures using ternary and quaternary III–V compounds deposited on a foreign semiconductor substrate, the maximum efficiency in space conditions, the best results in terms of power density, active lifetime, and minimal degradation over this period have been achieved. With the help of such solar cells, the efficiency range of 25-30% has been mastered. For a whole class of promising spacecraft, for example, large geostationary platforms, as well as spacecraft designed for transport operations in space using electric propulsion systems, the ability to fulfill modern targets is only possible using such highly efficient solar panels. Taking this into account, and also using many years of experience in designing GaAs-based solar cells, Kvant is developing work in this direction.

Creation of flexible thin-film solar cells based on amorphous silicon with maximum specific energy-mass characteristic and minimum cost.

This is a completely new direction in space photovoltaics. The most promising type of such photovoltaic converters at present are 3-stage solar cells based on amorphous silicon (a-Si). Originally designed for ground-based photovoltaics, amorphous silicon solar cells are currently being considered for use in space due to:

• the possibility of obtaining high energy-mass characteristics of solar cells is 4-5 times higher than that of solar cells made on the basis of single-crystal silicon, despite their lower initial efficiency;
• high radiation resistance;
• the possibility of reducing the unit cost of a solar battery by an order of magnitude and more compared to a single-crystal variant.

A significant advantage of flexible thin-film solar cells is their small starting (transport) volume, the possibility of creating easily deployable roll-type solar cells on their basis, etc.

As the basic technology for manufacturing photovoltaic converters based on amorphous silicon for space applications, the ground-based technology mastered by the Russian-American joint venture Sovlaks LLC (co-founders of NPP Kvant, ECD Ltd., USA) is considered. This technology ensures the formation of a cascade three-junction photovoltaic structure based on a-Si alloys on a thin ribbon substrate.

Modern projects of NPP Kvant in the field of space photovoltaics

• ISS: Russian segment of Zarya and Zvezda modules with solar converters with two-way sensitivity
• Large geostationary platforms "SiSat", "Express-A", "Express-AM", "KazSat", etc.
• Spacecraft for remote sensing of the Earth and meteorology "Monitor-E", "Meteor-3", etc.

The invention relates to electrical engineering, in particular to devices for generating electrical energy by converting light radiation into electrical energy, and can be used in the design and manufacture of small spacecraft with solar panels (SB). The technical result of the invention is: increasing the resistance of the SB to thermal shocks, to the effects of mechanical and thermomechanical loads, increasing the manufacturability of the design, increasing the active life of the SB of spacecraft, increasing functionality by expanding the temperature range of operation and optimizing the design of the SB, simplifying the switching system, which is achieved by increasing the strength of the connection of shunt diodes and solar cells, increasing the reproducibility of the manufacturing process of spacecraft SB by optimizing the manufacturing technology of shunt diodes and SB SB, as well as switching buses connecting the solar cell and shunt diodes, which are made of multilayer. The solar battery for small-sized space vehicles contains: panels with modules with solar cells (SC) glued to them, a shunt diode; switching buses connecting the front and back sides of the shunt diode with the SC, while the shunt diode is installed in a cutout in the corner of the SC, while the switching buses are made multilayer, consisting of molybdenum foil, on both sides of which a layer of vanadium or titanium, a layer of nickel and silver layer, respectively. 2 n. and 5 z.p. f-ly, 4 ill., 3 tab.

Drawings to the RF patent 2525633

Technical field

The invention relates to electrical engineering, in particular to devices for generating electrical energy by converting light radiation into electrical energy, and can be used in the design and manufacture of small spacecraft with solar panels (SB).

State of the art

The following requirements are imposed on the SB: maximum energy efficiency with a minimum mass, preservation of electrical and mechanical characteristics during storage, transportation on Earth and launch into the calculated orbit, long active life (SAS) in orbit with minimal degradation, which is expressed in power loss. In modern SBs, the SAS reaches 15 years, and requirements are being put forward to increase it to 20 years.

The main causes of degradation in orbit are violations of the structure of active elements, namely photoconverters (PC) and diodes under the influence of radiation, as well as violations resulting from the impact of temperature changes, thermal cycles. On different orbits, the range of temperature changes and the frequency of thermal cycles are different. For operating conditions in geostationary orbit, the upper temperature value is +100°C, the lower temperature is 170°C, the number of thermal cycles is 2000. In low orbits, the temperature range is smaller, the upper value is +100°C, the lower value is 100°C, but the number of thermal cycles in during the period of active existence in orbit is several tens of thousands.

It is known from the prior art (see N. S. Rauschenbach. The principles and technology of photovoltaic energy conversion. New York, 1980) that the SB consists of separate generators, including chains of solar cells (SC), inside the generators in anti-parallel with solar elements install shunt diodes. In addition to shunt diodes, to ensure reliable operation of the SB, diode protection is used, which is provided by blocking diodes.

In recent years, silicon solar cells have been replaced by more efficient multi-stage heterojunction solar cells based on A3B5 compounds grown on a germanium substrate (see P. R. Sharps, M. A. Stan, D. J. Aiken, B. Clevenger, J. S. Hill and N. S. Fatemi, High efficiency, multi-junction cells with monolithic bypass diodes, NASA/CP.2005-213431. Page 108-115). Each such SC is protected by a diode located with the SC in the same plane, and the diode has the same thickness as the SC. Usually in the SC, cuts are made at the corners, in which a triangular-shaped diode is placed (see US patents for inventions US 6353176, US 6034322 and US application for invention US 2008/0000523).

The prior art is known solar battery of spacecraft, located on a carbon fiber honeycomb panel. The bearing part of the honeycomb panel consists of two layers of carbon fiber, between which there is a honeycomb filler of aluminum foil. An electrical insulating film is glued onto the carbon-fiber surface intended for mounting the solar cell. The power generating part of the solar battery (modules) consists of solar cells connected in series or in series-parallel with the help of switching elements with thermomechanical compensators. A glass plate is glued to the front surface of each SC (see GLOBASTAR. Solar Generator Design 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 battery of space vehicles include low manufacturability of the design, a small temperature range of operation due to the low strength of soldered and welded joints of shunt diodes and solar cells. The high probability of damage to the interelement switching protruding above the front surface of the SB during its manufacture and routine maintenance, as well as the technological complexity of manufacturing the interelement switching, due to the need to place thermomechanical compensators in narrow interelement gaps, leads to a low resistance of the SB to the effects of thermal and mechanical loads .

The closest technical solution (prototype) in terms of technical essence and achieved effect is a solar battery of spacecraft, containing panels with modules glued to them, consisting of solar cells connected in series or in series-parallel using switching buses, where the switching tires are equipped with thermomechanical compensators, and to The front surface of each SE is glued with a protective glass plate, which is provided with additionally glued to the flat or curved surface of the frame with elastic elements having a given shape and size, where the internal volume of the elastic elements is filled with a sealant with the formation of a convex meniscus, and the SE is pressed against the elastic elements and fixed motionless, and switching busbars with thermomechanical compensators and shunt diodes are welded or soldered to the rear contacts of the ESS in areas free of sealant, and thermomechanical compensators are located between the back side of the ESS and the bearing surface of the frame in areas free of sealant (see Fig. patent of the Russian Federation for the invention RU 2250536).

The disadvantages of the well-known spacecraft solar battery include low manufacturability of the design, a small temperature range of operation due to the low strength of soldered and welded joints of shunt diodes and solar cells, and poor resistance of the SB to mechanical and thermomechanical loads. The molybdenum tire, which is 50 µm thick and has a multilayer special coating, is very rigid. When connecting switching busbars by welding, the electrical characteristics of the shunt diodes deteriorate, and in some cases, due to a rigid busbar, the weld point breaks out along with silicon, which leads to a low yield of suitable crystals after thermal cycling tests. At elevated temperatures, degradation of the SC occurs after soldering and welding, which leads to delamination of the contacts from the SC and, as a result, the exit from the working state of the SB cells.

From the prior art, a method is known for manufacturing solar cells of space vehicles with a shunt diode, including the manufacture of a solar cell based on a photovoltaic semiconductor substrate, the formation of shunt diodes on the front side of the solar cell, the connection of shunt diodes and solar cells of the solar cell of space vehicles, the connection using switching busbars of the solar cell (see US patent for invention US6635507).

To disadvantages known way low reproducibility of the manufacturing process due to the high probability of delamination (loss of adhesion) of the metallization on the working and non-working sides. In addition, when connecting the switching busbars by welding, the layers of the structure can be closed by the switching busbar, and the welding point is pulled out together with the substrate structure, which consequently leads to a low yield of good crystals after thermal cycling tests.

The closest technical solution (prototype) in terms of technical essence and achieved effect is a method for manufacturing spacecraft solar panels with an integrated shunt diode, including manufacturing a solar cell based on a photovoltaic semiconductor substrate with recesses for accommodating discrete shunt diodes, manufacturing discrete shunt diodes based on a semiconductor substrate, mounting discrete shunt diodes into recesses, contacting solar cells with shunt diodes using switching buses (see US patent for the invention US 5616185).

The disadvantages of the known manufacturing method include the low reproducibility of the manufacturing process due to the high probability of delamination (loss of adhesion) of the metallization during the formation of the metallization of the non-working side. In addition, when cutting into crystals, cracks form on silicon single-crystal substrates, and when connecting connecting busbars by welding, the welding point breaks out together with silicon, which leads, as a result, to a low yield of suitable crystals after thermal cycling tests (thermal shocks).

Disclosure of invention

The technical result of the claimed invention is:

Increasing the resistance of the SB to thermal shocks, to the effects of mechanical and thermomechanical loads, increasing the manufacturability of the design, increasing the life of the active existence of the SB of spacecraft, increasing the functionality by expanding the temperature range of operation and optimizing the design of the SB,

Simplification of the switching system, which is achieved by increasing the strength of the connection between shunt diodes and solar cells,

Increasing the reproducibility of the spacecraft SB manufacturing process by optimizing the manufacturing technology of shunt diodes and SB SB, as well as switching buses connecting SC and shunt 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:

shunt diode;

at the same time, the switching buses are made multilayer, consisting of molybdenum foil, on both sides of which a layer of vanadium or titanium, a layer of nickel and a layer of silver, respectively, are sequentially deposited.

In the 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 µm.

A method for manufacturing a solar battery for small spacecraft includes:

Connection of solar cells with shunt diodes using switching

at the same time, switching buses are made multilayer from molybdenum foil, on both sides of which a layer of vanadium or titanium, a layer of nickel and a layer of silver are sequentially applied, respectively.

In a preferred embodiment, a layer of vanadium or titanium, a layer of nickel and a layer of silver are applied sequentially from both sides to the prepared molybdenum foil by vacuum magnetron sputtering at a molybdenum foil temperature of 110-130 ° C with preliminary ion bombardment, and molybdenum foil with formed layers of vanadium or titanium, nickel and silver are annealed in vacuum at a temperature of 300-350°C.

Brief description of the drawings

The features and essence of the claimed invention are explained in the following detailed description, illustrated by drawings, where the following is shown.

Figure 1 shows the solar cell installed on the side with the help of switching tires shunt diode.

Figure 2 schematically shows the layered structure of the switching

Figure 3 shows the algorithm of the manufacturing method of the Sat spacecraft.

Figure 4 shows the values ​​of internal mechanical stresses calculated from experimentally measured deformations in the metal layers of the switching tires formed at different temperatures of the molybdenum foil.

In figure 4, the graphs in parentheses indicate the optimal operating temperature range of molybdenum foil during sputtering. Figure 1 indicates the following:

1 - shunt diode;

2 - switching bus connecting the front side of the shunt diode (1) with the solar cell (4);

3 - switching bus connecting the reverse side of the shunt diode (1) with the solar cell (4);

4 - solar cell (SC);

Figure 2 indicates the following:

5 - prepared molybdenum foil;

6 - layer of vanadium or titanium;

7 - nickel layer;

8 - a layer of silver.

Implementation and exemplary implementation of the invention

The claimed method was used in the implementation of a group technology for the manufacture of solar panels for spacecraft and consists of the following sequence of technological operations (see figure 3): manufacture of solar cells based on a photovoltaic semiconductor substrate, manufacture of shunt diodes based on a photovoltaic semiconductor substrate, manufacture of switching tires , which includes the preparation of molybdenum foil and the metallization of the prepared molybdenum foil by vacuum magnetron sputtering on both sides with layers of vanadium, nickel and silver at a molybdenum foil temperature of 110-130 ° C with preliminary ion bombardment, then the molybdenum foil is annealed with formed layers of vanadium or titanium , nickel and silver in vacuum at a temperature of 300-350°C, welding of switching buses to shunt diodes, testing shunt diodes for thermal cycling and thermal shock, connecting solar cells with shunt diodes using switching tires and performing output control of the solar battery of spacecraft.

The thickness of the molybdenum foil was chosen based on the highest tear-off force of the welded switching busbar to the front and back sides of the shunt diode after thermal shock testing.

The pull-off force of the welded switching busbar from the shunt diode was determined as follows: the molybdenum foil was prepared in several stages, after which the molybdenum foil was thinned to the following thicknesses: 6±0.1 µm, 7.5±0.1 µm, 10±0.1 µm, 13±0.1 µm. Then, layers of vanadium, nickel and silver were deposited on the prepared molybdenum foil by vacuum magnetron sputtering on both sides at a molybdenum foil temperature of 110-130°C with preliminary ion bombardment.

After that, the molybdenum foil with the formed layers of vanadium or titanium, nickel and silver was annealed in vacuum at a temperature of 300-350°C and punched out from the molybdenum foil of the switching buses. After that, control welding of the switching tires to the front and back sides of the shunt diodes and control of the force of separation of the switching tires from the shunt diodes were carried out (see Table 1).

Then the thermal shock tests of the welded switching busbars to the shunt diodes were carried out, which consisted in conducting 450 thermal shock cycles from a temperature of -180°C (liquid nitrogen vapor) to 120°C on specialized equipment. After that, the electrical parameters of the shunt diodes were measured, which showed a slight increase in the forward voltage against the background of unchanged leakage currents and reverse voltage. Then, the force of separation of the switching tires from the shunt diodes was monitored (see Table 2).

As a result of the tests, an increase in the separation force was revealed for all variants of the thickness of the switching tires from the shunt diodes with a slight change in the electrical characteristics of the shunt diodes. Based on Table 2, it was found that the optimal thickness of the molybdenum foil is 10 ± 0.1 μm, since the maximum pull-off force of the busbar from the shunt diode is provided.

The temperature of the molybdenum foil during the technological operation of deposition of metals was chosen based on the minimum stresses in the resulting structure (see figure 4). Internal stresses were determined as follows: single-cantilever microbeams were formed by magnetron sputtering of V-Ni-Ag metal films on prepared molybdenum foil with photolithography and plasma-chemical etching of metals. The obtained samples of single-cantilever microbeams were examined using an optical microscope Axio Imager (Carl Zeiss) at a magnification of 6000x. The dimensions of the beam structure and the direction of deformation were measured. The form of deformation was determined by the deviation of microbeams at different points of its length from the surface. After that, using mathematical processing according to the Stoney formula, the stresses of the beams were calculated. The curvature of the beam was found by measuring the deviation of the shank of a single-cantilever microbeam. These modes were chosen based on considerations of reproducibility technological process, which is provided if, when connecting the switching buses by welding, the welding point does not break out (see Table 3).

According to the proposed design and manufacturing method, solar panels were made for small-sized spacecraft, including packageless shunt triangular-shaped diodes with a reverse voltage of 100 V and a direct current of 2 A and cascade photoconverters based on A 3 V 5 connections.

Prior to the use of the claimed technical solution, silver switching buses were used, which were welded to shunt diodes and solar cells. Testing of the diodes showed low resistance to thermal shocks (the structure was destroyed after 10-15 thermal shocks from -180°C to +100°C), and the percentage of yield of good diodes in terms of electrical characteristics at the thermal cycling stage was no more than 70% of the good diodes after assembly, and in the remaining 30%, the structure was destroyed in the welding zone (interlayer destruction along the base materials when exposed to elevated and low temperatures) during the control of the strength of the welded joint. The detachment force of the metallization from the crystal was 50-100 g/mm 2 , and after using this technical solution it was more than 150 g/mm 2 , as a result of which the percentage of yield of good diodes at the thermal cycling stage increased to 85%.

CLAIM

1. Solar battery for small spacecraft contains:

Panels with modules with solar cells (SC) glued to them,

shunt diode;

Switching busbars welded to the front and back sides of the shunt diodes and connecting the front and back sides of the shunt diode to the solar cell, while the shunt diode is installed in a cutout in the corner of the solar cell,

characterized in that

switching buses are made multilayer, consisting of molybdenum foil, on both sides of which a layer of vanadium or titanium, a layer of nickel and a layer of silver are successively deposited, respectively.

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 microns.

5. A method for manufacturing a solar battery for small spacecraft, including:

Production of solar cells (SC) based on a photovoltaic semiconductor substrate with a cutout in the corner for shunt diodes,

Fabrication of shunt diodes based on a photovoltaic semiconductor substrate,

Manufacturing of switching busbars,

Welding of switching buses to the front and back sides of the shunt diodes,

Installation of shunt diodes in the cutout in the corner of the SE,

Connection of solar cells with shunt diodes using switching buses,

characterized in that

switching buses are made multi-layered from molybdenum foil, on both sides of which a layer of vanadium or titanium, a layer of nickel and a layer of silver are successively deposited, respectively.

6. The method according to claim 5, characterized in that a layer of vanadium or titanium, a nickel layer and a silver layer are applied sequentially from both sides to the prepared molybdenum foil by vacuum magnetron sputtering at a molybdenum foil temperature of 110-130°C with preliminary ion bombardment.

7. The method according to claim 6, characterized in that the molybdenum foil with the formed layers of vanadium or titanium, nickel and silver is annealed in vacuum at a temperature of 300-350°C.

These are photovoltaic converters - semiconductor devices that convert solar energy into direct electricity. Simply put, these are the main elements of the device that we call "solar panels". With the help of such batteries in space orbits artificial satellites Earth. Such batteries are made here in Krasnodar - at the Saturn plant. The plant management invited the author of this blog to look at the production process and write about it in his diary.

1. The enterprise in Krasnodar is part of the structure of the Federal Space Agency, but Saturn is owned by the Ochakovo company, which literally saved this production in the 90s. The owners of Ochakovo bought out a controlling stake, which almost went to the Americans. Ochakovo invested heavily here, purchased modern equipment, managed to retain specialists, and now Saturn is one of the two leaders in the Russian market for the production of solar and storage batteries for the needs of the space industry - civil and military. All the profit that Saturn receives remains here in Krasnodar and goes to the development of the production base.

2. So, it all starts here - on the site of the so-called. gas phase epitaxy. There is a gas reactor in this room, in which a crystalline layer is grown on a germanium substrate for three hours, which will serve as the basis for a future photocell. The cost of such an installation is about three million euros.

3. After that, the substrate still has a long way to go: electrical contacts will be applied to both sides of the photocell (moreover, on the working side, the contact will have a “comb pattern”, the dimensions of which are carefully calculated to ensure maximum passage of sunlight), an anti-reflective coating will appear on the substrate coating, etc. - in total more than two dozen technological operations at various installations before the photocell becomes the basis of a solar battery.

4. Here, for example, is the installation of photolithography. Here, on the photocells, “patterns” of electrical contacts are formed. The machine performs all operations automatically, according to a given program. Here, the light is appropriate, which does not harm the light-sensitive layer of the photocell - as before, in the era of analog photography, we used "red" lamps.

5. In the vacuum of the sputtering installation, electrical contacts and dielectrics are applied using an electron beam, as well as antireflection coatings are applied (they increase the current generated by the photocell by 30%).

6. Well, the photocell is ready and you can start assembling the solar battery. Tires are soldered to the surface of the photocell in order to then connect them to each other, and a protective glass is glued on them, without which in space, under radiation conditions, the photocell may not withstand loads. And, although the thickness of the glass is only 0.12 mm, a battery with such photocells will work for a long time in orbit (more than fifteen years in high orbits).

6a

6b

7. The electrical connection of the photocells with each other is carried out by silver contacts (they are called shank) with a thickness of only 0.02 mm.

8. To obtain the desired voltage in the network, produced by the solar battery, the photocells are connected in series. This is how a section of series-connected photocells looks like (photoelectric converters - that's right).

9. Finally, the solar panel is assembled. Only part of the battery is shown here - the panel in layout format. There can be up to eight such panels on the satellite, depending on how much power is needed. On modern communication satellites, it reaches 10 kW. Such panels will be mounted on a satellite, they will open in space like wings and with their help we will watch satellite TV, use satellite Internet, navigation systems (Glonass satellites use Krasnodar solar panels).

9a

10. When the spacecraft is illuminated by the Sun, the electricity generated by the solar battery feeds the systems of the apparatus, and the excess energy is stored in the battery. When the spacecraft is in the shadow of the Earth, the spacecraft uses the electricity stored in the battery. The nickel-hydrogen battery, having a high energy capacity (60 Wh/kg) and an almost inexhaustible resource, is widely used in spacecraft. The production of such batteries is another part of the work of the Saturn plant.

In this picture, Anatoly Dmitrievich Panin, holder of the medal of the Order of Merit for the Fatherland, II degree, is assembling a nickel-hydrogen battery.

10a

11. Assembly site for nickel-hydrogen batteries. The filling of the battery is being prepared for placement in the case. The filling is positive and negative electrodes separated by separator paper - in them the transformation and accumulation of energy takes place.

12. Installation for electron-beam welding in vacuum, with which the battery case is made of thin metal.

13. A section of the workshop where the cases and parts of the batteries are tested for the effects of high pressure.
Due to the fact that the accumulation of energy in the battery is accompanied by the formation of hydrogen, and the pressure inside the battery increases, leak testing is an integral part of the battery manufacturing process.

14. The case of a nickel-hydrogen battery is a very important part of the entire device operating in space. The body is designed for a pressure of 60 kg·s/cm 2 , during testing the rupture occurred at a pressure of 148 kg·s/cm 2 .

15. Batteries tested for strength are filled with electrolyte and hydrogen, after which they are ready for use.

16. The body of the nickel-hydrogen battery is made of a special alloy of metals and must be mechanically strong, light and have high thermal conductivity. Batteries are installed in cells and do not touch each other.

17. Accumulators and batteries assembled from them are subjected to electrical tests at our own production facilities. In space, it will be impossible to fix or replace anything, so every product is carefully tested here.

17a

17b

18. All space technology is subjected to tests for mechanical effects using vibration stands that simulate the load during the launch of the spacecraft into orbit.

18a

19. In general, the Saturn plant made the most favorable impression. The production is well organized, the workshops are clean and bright, the people are qualified, it is a pleasure and very interesting to communicate with such specialists for a person who is at least to some extent interested in our space. Left the Saturn good mood- It's always nice to see a place where they don't engage in empty chatter and don't shift papers, but do a real, serious business, successfully compete with the same manufacturers in other countries. There would be more of this in Russia.

Photos: © drugoi

P.S. Blog of the Vice President for Marketing of the Ochakovo company

The invention relates to rocket and space technology, and in particular to structural elements of solar panels of spacecraft. The bearing panel of the solar battery of the spacecraft contains a frame and bearing upper and lower bases. Between the mentioned bases and the frame, a filler in the form of honeycombs is hermetically installed and load-bearing partitions are perpendicular to the bases. To communicate the internal volumes of the honeycombs with each other, each of the variants of the invention provides for the implementation of drainage holes in the side surfaces of each honeycomb of the filler and power partitions. To communicate the internal volumes of the honeycombs with the external environment, the first variant of the invention provides for the implementation of drainage holes in at least one frame element, the second variant of the invention provides for the implementation of drainage holes in the lower base of the panel evenly over its surface area, and the third variant of the invention provides for the implementation of drainage holes at least least in one frame element and in the lower base of the panel evenly over its surface area. At the same time, the total areas of drainage holes in the said structural elements of the carrier panel are determined taking into account the total volume of the gaseous medium in the honeycombs, the discharge coefficients of the drainage holes, and the maximum pressure drop of the gaseous medium along the flight path of the launch vehicle acting on the bases of the panel. EFFECT: invention makes it possible to increase the structural strength of spacecraft solar panels without increasing their weight, to simplify the manufacturing and installation technology of the panels, and to increase the reliability of their operation. 3 n.p. f-ly, 4 ill.

The invention relates to the field of aerodynamics aircraft(LA) and can be used in rocket science in the design and creation of solar panels (SB) of spacecraft (SC) made according to a three-layer carrier scheme.

Known and widely used in aviation in the manufacture of aircraft elements (fuselage, empennage, wing, etc.) are panels made according to a three-layer carrier scheme, containing a frame (frame) carrying the upper and lower bases, between which a honeycomb filler is installed.

Designed for the perception and transmission of distributed loads acting on aircraft elements, panels made according to a three-layer scheme with honeycomb filler provide greater rigidity and high load-bearing capacity. When the panel is loaded, the shear-hard and lightweight honeycomb filler perceives transverse shear and protects thin carrier layers from buckling under longitudinal compression.

The disadvantages of this technical solution include the increased weight of the frame elements and the supporting bases of the panels due to significant pressure drops acting on the panel elements along the flight path of the aircraft when the flight altitude of the aircraft changes.

Known used in rocket science panel SC spacecraft, intended for installation on them sensitive elements (photoelectric converters) of the spacecraft power supply system. The panels are also made according to a three-layer bearing scheme and contain a frame, bearing the upper and lower bases, between which the filler in the form of honeycombs is hermetically installed, as well as load-bearing partitions, hermetically installed perpendicular to the bases to increase the rigidity of the panel. To reduce the weight of the construction of SB panels, the frame, load-bearing bases and partitions are made of lightweight materials.

SC load-bearing panels used in rocket science, as well as panels used in aviation, provide greater rigidity and high load-bearing capacity of the three-layer structure of the SB panel with honeycomb core.

The disadvantages of this technical solution include the reduced structural strength of the SB load-bearing panels and the possibility of losing its general and local stability in case of deviations in the panel manufacturing and operation technology, due to more significant aerogasdynamic loads acting on the elements of the SB SC panels, compared with aviation loads. At the same time, the external pressure acting on the SC SB panel along the flight path of the launch vehicle (LV) varies over a wider range: from atmospheric (at the Earth level at the launch of the LV) to almost zero when it is launched into interplanetary space, and the pressure inside the sealed panel along the flight path, the launch vehicle remains atmospheric.

The objective of the invention is to increase the structural strength of the load-bearing panels of the spacecraft SB without increasing their mass when the spacecraft is launched by a launch vehicle into interplanetary space.

The problem is solved in such a way (option 1) that in the carrier panel of the SB KA, containing a frame, bearing upper and lower bases, between which the filler in the form of honeycombs is hermetically installed, power partitions, hermetically installed perpendicular to the bases, according to the invention in the side surfaces of each honeycomb of the filler and partitions, through drainage holes are made, communicating the internal volumes of the honeycombs with each other, and in the frame, at least in one frame element, drainage holes are made, communicating the internal volumes of the cells with the external environment, while the total effective area of ​​the drainage holes in the honeycombs, partitions and frame is determined from the ratios:

S 2 [cm 2 ] - the total area of ​​the drainage holes in the frame;

a, b are coefficients depending on the parameters of the launch vehicle trajectory, approximating the curve of dependence of the effective area of ​​the drainage holes in the frame on the maximum pressure drop along the trajectory acting on the bases of the panels.

The problem is also solved in such a way (option 2) that in the carrier panel of the SA SC, containing a frame, bearing upper and lower bases, between which the filler in the form of honeycombs is hermetically installed, power partitions, hermetically installed perpendicular to the bases, according to the invention in the side surfaces of each honeycomb filler and partitions, drainage holes are made that communicate the internal volumes of the honeycombs to each other, and in the lower base of the panel, drainage holes are made uniformly over its surface area, communicating the internal volumes of the honeycombs with the external environment, while the total effective area of ​​the drainage holes in the honeycombs, partitions and the lower base determined from the ratios:

S 1 [cm 2 ] - the total area of ​​the drainage holes in the end surface of the cells;

S 3 [cm 2 ] - the total area of ​​the drainage holes in the lower base;

V [m 3 ] - the total volume of the gaseous medium in honeycombs;

μ.GIF; 1 - flow rate of drainage holes in honeycombs and partitions;

μ.GIF; 3 - flow rate of drainage holes in the lower base;

∆.GIF; P [kgf/cm 2 ] - the maximum pressure drop of the gaseous medium along the flight path of the launch vehicle, acting on the base of the panel;

a, b are coefficients depending on the parameters of the launch vehicle trajectory, approximating the curve of dependence of the effective area of ​​drainage holes in the bases of the panels on the maximum pressure drop along the trajectory acting on the bases of the panel.

The problem is also solved in such a way (option 3) that in the supporting panel of the SA SC, containing a frame, bearing upper and lower bases, between which the filler in the form of honeycombs is hermetically installed, power partitions, hermetically installed perpendicular to the bases, according to the invention in the side surfaces of each honeycomb the filler and partitions are made through drainage holes that communicate the internal volumes of the honeycombs with each other, and in the frame, at least in one frame element, and in the lower base of the panel, drainage holes are made uniformly over its surface area, communicating the internal volumes of the honeycombs with the external environment, with In this case, the total effective area of ​​drainage holes in the honeycombs, partitions, frame and lower base is determined from the ratios:

S 1 [cm 2 ] - the total area of ​​the drainage holes in the end surface of the cells;

S 2 , S 3 [cm 2 ] - the total area of ​​the drainage holes in the frame and bottom base, respectively;

V [m 3 ] - the total volume of the gaseous medium in honeycombs;

μ.GIF; 1 - flow rate of drainage holes in honeycombs and partitions;

μ.GIF; 2, μ.GIF; 3 - flow coefficient of drainage holes in the frame and bottom base of the panel, respectively;

∆.GIF; P [kgf/cm 2 ] - maximum differential pressure of the gaseous medium along the flight path of the launch vehicle, acting on the base of the panel;

The technical results of the invention are:

Reduction of pressure drops acting on the bases and sensitive elements of the SB panel at the minimum allowable pressure drops acting on the walls of the filler honeycombs;

Determination of the effective area of ​​drainage holes in the honeycomb, frame, load-bearing bases and panel partitions;

Determination of the influence of the trajectory parameters (M number, flight altitude H) on the effective area of ​​the drainage holes.

The essence of the invention is illustrated by the diagrams of the SB KA panel and the graph of changes in excess pressures acting on its elements.

Figures 1, 2 and 3 show the diagrams of the SA panel of the spacecraft, made respectively in options 1, 2 and 3, and its fragments are highlighted, where:

2 - upper base;

3 - lower base;

4 - filler;

5 - partitions;

6 - drainage holes;

7 - sensitive elements.

Here, the arrows show the direction of the flow of the gas medium in the honeycombs of the panel filler and its outflow into the external environment.

Figure 4 shows the dependence of the maximum along the flight path of the LV pressure drop Δ.GIF; Р(Δ.GIF; Р=Рvn-Рnar) of the gaseous medium acting on the bases of the panels, from the relative effective area of ​​the flow sections of the drainage holes μ.GIF; S/V, where:

Pvn - pressure of the gaseous medium inside the panel (in the honeycombs of the filler);

Рnar - pressure of the gaseous medium outside the panel.

Carrier panel SB KA (figure 1, 2, 3) contains a frame 1, bearing the upper base 2 and the lower base 3, as well as power partitions 5 installed perpendicular to these bases. Filler 4 in the form of honeycombs is hermetically installed between the bases. On the upper base 2, sensitive elements 7 of the spacecraft power supply system are installed.

In the side surfaces of each honeycomb filler 4 and power partitions 5, in contrast to the prototype, in each embodiment, drainage holes 6 are made, communicating the internal volumes of the honeycombs with each other and with the external environment (see view A and section BB).

In option 1 (figure 1) the internal volumes of cells communicate with the external environment through drainage holes 6 made in the frame 1, at least in one of its elements.

In option 2 (figure 2) the internal volumes of honeycombs communicate with the external environment through drainage holes 6 made in the lower base 3, evenly spaced over the area of ​​its base.

In option 3 (figure 3), the internal volumes of the cells communicate with the external environment through drainage holes 6 made in the frame 1, at least in one of its elements, as well as in the lower base 3, evenly spaced over the area of ​​its base.

Due to the uniform arrangement of drainage holes over the area of ​​the panel bases, a uniform or close to uniform distribution of pressure in the core cells and, consequently, pressure drops acting on the panel bases is ensured. This eliminates stress concentrations at the junction of the panel elements from uneven pressure drops, which leads to a simplification of the technology for manufacturing panels and an increase in the reliability of its operation in the presence of hidden defects in its manufacture, for example, when individual elements of the core honeycombs are not glued to bearing bases.

The choice of panel drainage option is determined by the allowable operational loads acting on the bases of the panels along the flight path of the launch vehicle, taking into account the design and technological features of the panels manufacturing.

The total effective area of ​​the drainage holes in the frame 1, in the filler honeycombs 4, partitions 5 and lower base 3 for a given launch vehicle flight path is determined by relations (1), (2) and (3), for options 1, 2 and 3, respectively, with taking into account the coefficients a, b included in these relations, depending on the parameters of the launch vehicle trajectory.

Formulas (1), (2) and (3) contain mathematical description dependences of the relative total effective area of ​​drainage holes μ.GIF; ·S/V from the maximum pressure difference along the flight path PH Δ.GIF; P and obtained from the analysis of the flow of the gaseous medium in the system of gas-dynamic interconnected tanks formed by drained honeycombs of the filler 4 with power partitions 5, the upper base 2 and the lower base 3, followed by its outflow into the external environment.

In rocket science, the frame 1 is made of carbon fiber, the supporting bases 2 and 3, as well as the power partitions 5, are made of titanium. The filler 4 in the form of a honeycomb is made of aluminum alloy and is hermetically fixed to the upper base 2 and the lower base 3 of the panel using, for example, VKV-9 aviation glue. Also, sensitive elements 7 SB are attached to the upper base 2.

Carrier panel SAT KA works as follows.

Since in the side surfaces of each cell core 4 and panel elements (figure 1, 2 and 3), unlike the prototype, drainage holes 6 are made, during the flight of the spacecraft as part of the head unit of the launch vehicle, as well as in the autonomous flight of the spacecraft, after resetting the fairings head block, the gaseous medium flows between the cells of the filler 4, power partitions 5 and flows through the drainage holes in the frame 1 and the lower base 6 into the external environment (see section on BB). The overflow of the gaseous medium occurs with an insignificant delay in equalizing the pressure in the cells of the filler 4.

In this case, the outflow of the gaseous medium from the honeycombs of the filler 4 into the external environment occurs at subsonic speed with its non-locking in the honeycombs of the filler 4, since the total effective areas μ.GIF; 2 ·S 2 drainage holes 6 in frame 1 and μ.GIF; 3 ·S 3 - in the lower base 3 are made greater than or equal to the total effective area μ.GIF; 1 S 1 in the honeycomb filler 4 with power partitions 5 (μ.GIF; 2 S 2 ≥.GIF; μ.GIF; 1 S 1 , μ.GIF; 3 S 3 ≥.GIF; μ.GIF; 1 S 1).

During the flight of the spacecraft as part of the head unit of the launch vehicle, the maximum pressure drop Δ.GIF is realized; P (figure 4), acting on the base panels 2 and 3, in accordance with the formulas (1), (2) and (3). In this case, the gaseous medium from the filler cells 4 flows into a closed volume under the head fairing, the maximum allowable pressure drop in which, compared with the outer one along the LV flight path, is determined according to a well-known technical solution using the compartment drainage system.

In an autonomous flight of a spacecraft, an internal pressure Р ВН is established inside the body panel, which is close to atmospheric (static ambient atmosphere). Differences Δ.GIF; P pressure in this case between the honeycombs of the filler 4, as well as the internal pressure Рvn in the honeycombs of the filler 4 and the external environment Рnar, acting on the upper base 2 and the lower base 3 of the panel, are close to zero.

Thus, the pressure drops acting on the elements of the panels and the sensitive elements of the spacecraft power supply system installed on it are reduced. Thus, the structural strength of the SB spacecraft is increased without increasing the mass of the spacecraft, which leads to the fulfillment of the task.

In addition, due to the reduction of pressure drops acting on the elements of the panels, the manufacturing and installation technology of the SB KA panel is simplified and the reliability of its operation is increased.

Calculations carried out for the hull panel developed for the Yamal spacecraft, launched by the Proton launch vehicle, showed that the pressure drops Δ.GIF; P, acting on the basis of the panel, in comparison with the prototype, are reduced by an order of magnitude and almost approach zero.

At present, the technical solution has passed experimental testing and is being implemented on spacecraft being developed by the enterprise.

The technical solution can be used for various types of spacecraft: near-Earth, interplanetary, automatic, manned and other spacecraft.

The technical solution can also be applied in aviation, for example, when using the SB panel as part of an aircraft wing element. In this case, the effective area of ​​the drainage holes in the panel elements is determined taking into account the maximum pressure drops acting on the wing elements along the aircraft flight path.

Literature

1. Aviation. Encyclopedia. M.: TsAGI, 1994, p. 529.

2. At the turn of two centuries (1996-2001). Ed. acad. Yu.P.Semenova. M.: RSC Energia named after S.P. Korolev, 2001, p. 834.

3. Patent RU 2145563 C1.

Claim

1. The carrier panel of the solar battery of the spacecraft, containing a frame, bearing the upper and lower bases, between which the filler in the form of honeycombs is hermetically installed and power partitions perpendicular to the bases, characterized in that through drainage holes are made in the side surfaces of each honeycomb of the filler and power partitions, communicating the internal volumes of the honeycombs with each other, and in at least one frame element there are drainage holes that communicate the internal volumes of the honeycombs with the external environment, while the total effective area of ​​the drainage holes in the honeycombs, load-bearing partitions and the frame is determined from the ratios

S 2 - the total area of ​​the drainage holes in the frame, cm 2;

μ.GIF; 2 - flow rate of drainage holes in the frame;

a, b - dependent on the parameters of the trajectory of the launch vehicle, the coefficients approximating the curve of dependence of the effective area of ​​the drainage holes in the frame on the maximum pressure drop along the trajectory acting on the base of the panel.

2. The carrier panel of the solar battery of the spacecraft, containing a frame, bearing the upper and lower bases, between which the filler in the form of honeycombs is hermetically installed and power partitions perpendicular to the bases, characterized in that drainage holes are made in the side surfaces of each honeycomb of the filler and power partitions, communicating the internal volumes of the honeycombs are interconnected, and in the lower base of the panel, drainage holes are made uniformly over its surface area, communicating the internal volumes of the honeycombs with the external environment, while the total effective area of ​​the drainage holes in the honeycombs, load-bearing partitions and the lower base of the panel is determined from the ratios

μ.GIF; 1 S 1 /V=a Δ.GIF; P-b,

where S 1 - the total area of ​​the drainage holes in the side surfaces of honeycombs and power partitions, cm 2;

S 3 - the total area of ​​the drainage holes in the lower base of the panel, cm 2;

V is the total volume of the gaseous medium in honeycombs, m 3 ;

μ.GIF; 1 - flow rate of drainage holes in the side surfaces of honeycombs and power partitions;

μ.GIF; 3 - flow rate of drainage holes in the lower base of the panel;

∆.GIF; P is the maximum pressure drop of the gaseous medium along the flight path of the launch vehicle, acting on the base of the panel, kgf/cm 2 ;

a, b are coefficients depending on the parameters of the launch vehicle trajectory, approximating the curve of dependence of the effective area of ​​drainage holes in the lower base of the panel on the maximum pressure drop along the trajectory acting on the bases of the panel.

3. The carrier panel of the solar battery of the spacecraft, containing a frame, bearing the upper and lower bases, between which the filler in the form of honeycombs is hermetically installed and power partitions perpendicular to the bases, characterized in that through drainage holes are made in the side surfaces of each honeycomb of the filler and power partitions, communicating the internal volumes of the honeycombs with each other, and in at least one frame element and in the lower base of the panel, drainage holes are made uniformly over its surface area, communicating the internal volumes of the honeycombs with the external environment, while the total effective area of ​​the drainage holes in the honeycombs, power partitions, frame and lower base of the panel is determined from the ratios

μ.GIF; 1 S 1 /V=a Δ.GIF; P-b,

μ.GIF; 2 S 2 /V≥.GIF; μ.GIF; 1 S 1 /V,

μ.GIF; 3 S 3 /V≥.GIF; μ.GIF; 1 S 1 /V,

where S 1 - the total area of ​​the drainage holes in the side surfaces of honeycombs and power partitions, cm 2;

S 2 , S 3 - total areas of drainage holes in the frame and bottom base of the panel, respectively, cm 2 ;

V is the total volume of the gaseous medium in honeycombs, m 3 ;

μ.GIF; 1 - flow rate of drainage holes in the side surfaces of honeycombs and power partitions;

μ.GIF; 2, μ.GIF; 3 - flow coefficients of drainage holes in the frame and bottom base of the panel, respectively;

∆.GIF; P is the maximum pressure drop of the gaseous medium along the flight path of the launch vehicle, acting on the base of the panel, kgf/cm 2 ;

a, b are coefficients depending on the parameters of the launch vehicle trajectory, approximating the curve of dependence of the effective area of ​​the drainage holes in the frame and the lower base of the panel on the maximum pressure drop along the trajectory acting on the bases of the panel.