RU2555891C1 - Microstructural multilayered screen and vacuum isolation of spacecrafts - Google Patents

Abstract

FIELD: aviation.

SUBSTANCE: invention relates to the multilayered screen and vacuum isolation (SVI) with microstructural elements for spacecrafts. Each layer of SVI is implemented as a substrate on which the heat reflecting elements designed as massif of rectangular microplates are fixed. Each microplate is fixed on a substrate with a gap 10…20 mcm. On the substrate side facing towards the spacecraft the flutes with rectangular or trapezoidal section, and also longitudinal depressions with semicircular cross-section are implemented. The second and the subsequent layers of SVI are attached to the previous layers through the spherical spacers installed between plates. The diameter of spacers is no less than the size of the named gap. In the places of installation of spacers the layers of silicon dioxide with the thickness 0.5…1 are applied. The aluminium coating with the thickness 0.1…0.3 microns with the reflection factor 0.7-0.9 is applied on external surface of microplates and open surfaces of the substrate. Microplates can be bimorph. Microplates are manufactured with conducting buses on the surface of the silicon substrate.

EFFECT: decrease of weight and overall dimensions of EVI and spacecrafts.

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Description

Technical field

The invention relates to microstructural devices in the field of astronautics and can be used as a separate screen-vacuum insulation (EVI) of spacecraft (SC), in particular, nano- and picosatellites, and also as an element of a complex system of thermal insulation of the SC or its components.

State of the art

The prior art active vacuum insulation, consisting of a chamber closed by two layers of metal, glass-like spacers located in the chamber between the side walls and providing a gap between them (US patent No. US 5562154, publ. 8.10.1996). A high vacuum has been created in the chamber. The composition of the insulation includes equipment for regulating the coefficient of thermal insulation of the surface. The first version of the on / off apparatus and methodology includes a metal hydride to produce hydrogen in the chamber in case of overheating, and a tube between the metal hydride and the chamber to allow the return of hydrogen to the metal hydride. A second type of on / off apparatus and methodology includes a coating on a metal surface with a varying emissivity, in which the emissivity changes under the influence of heat or electricity. In addition, the second type of on / off equipment and methodology includes metal-to-metal contact devices that can be activated to close / open heated tracks or thermal circuits between the metal side walls.

The disadvantages of the known technical solution is the large weight of the insulation for its installation on small spacecraft, the difficulties with its installation on the spacecraft with a large number of conclusions and equipment on the outer casing, which requires the creation of complex insulation. In addition, the use of metallic hydride in the system to produce hydrogen limits the life of the insulation, and the use of hydrogen requires additional precautions. The known design has a low speed due to its significant heat capacity, which negates the benefits of active control when installing insulation on unstabilized spacecraft.

The prior art insulation of space technology, used to prevent heat transfer between system elements and the surrounding space, occurring at very low temperatures and pressures below 10 Torr, is a barrier to heat transfer, consisting of many layers of a thin metallized plastic film with a metallization thickness of less than 1 microinch and an emissivity of less than 0.06 (US patent No. US 3,244,224, publ. 5.04.1966). The specified thin metal coating and the material of the plastic film of each layer provide low thermal conductivity through the side wall. The layers are located face to face relative to each other and are held, as a rule, through point contact between adjacent layers. The set of layers includes at least two pairs of layers, with two layers of a pair connected together at a point, and the second layer is tightly stretched on the bearing surface.

The disadvantages of the known technical solution is the large thickness of the structure, the need to use a large number of layers (100 layers or more), which implies the high complexity of the manufacture of insulation, the large weight of the system and the difficulty of installing insulation on the surface of a spacecraft having a large number of external elements.

The prior art screen-vacuum thermal insulation of a spacecraft with an external combined coating, consisting of a polymer substrate, an electrically conductive layer with a wear-resistant layer on the outer surface and a reflective layer on the inner surface (patent for the invention of the Russian Federation No. RU 2397926, publ. 08.27.2010). The insulation contains a temporary protective layer on the outer surface and a reinforcing polymer network on the inner surface. EFFECT: increase of reliability, efficiency, reduction of weight and dust separation of screen-vacuum thermal insulation.

The disadvantages of this technical solution include a rather high thermal conductivity due to the absence of vacuum gaps, large dimensions and weight of thermal insulation due to the use of a large number of different layers, the need to use a temporary protective coating to protect against damage and contamination of the thermal insulation, difficulties with installing the thermal insulation on the surface small spacecraft with a large number of external elements.

The closest in technical essence is the temperature-controlled coating and the method of its installation on a spacecraft containing a substrate in the form of individual elements of a transparent radiation-resistant material, the back surface of which is covered with reflective, protective and adhesive layers, and on the outer surface there is a transparent electrically conductive coating covering the end the surface of the substrate and in contact with the protective layer (patent for the invention of the Russian Federation No. RU 2356074, publ. 05.20.2009). The thickness of the individual elements is not less than 0.08 mm, the electrical resistance of the electrically conductive coating is not less than 2 and not more than 1 × 10 ⁵ kOhm / m ² . During installation, an electrically conductive adhesive layer is applied to the surface of the spacecraft, after which an auxiliary tape with an adhesive layer with individual elements glued is pressed to the surface of the spacecraft to remove air bubbles and excess glue. After curing the adhesive layer, the auxiliary tape is peeled off. The material of the adhesive layer is low viscosity in the initial state and elastic after curing, and contains a corrosion inhibitor, as well as a fibrous conductive material, the length of the fibers of which is greater than the thickness of the adhesive layer.

The main disadvantages of the prototype are:

- low efficiency of thermal insulation due to the lack of vacuum gaps;

- a large mass of the structure due to the overall dimensions and the materials used;

- high thermal conductivity between the adhesive layer and the spacecraft body due to the large contact area;

- the need to use an auxiliary tape with a sticky layer in compliance with a given gap between the end surfaces of the elements, which complicates the design and technology of manufacture and assembly of the coating;

- the need to use a protective layer between the reflective and adhesive layers, which complicates the design and manufacturing technology of the coating.

Disclosure of invention

The technical result of the invention is to reduce the weight and overall dimensions of the EVI KA.

The technical result is achieved by the fact that in the microstructural multilayer EVI of spacecraft, each layer is made in the form of a substrate mounted on the surface of the apparatus. Heat-reflecting elements are fixed on the substrate. The heat-reflecting elements are made in the form of an array of rectangular microplates. Each rectangular microplate is mounted on a substrate with a gap of 10 .... 20 μm from the surface of the substrate. On the substrate surface facing the device to be protected, grooves of a rectangular or trapezoidal section are made. The second and subsequent EVI layers are attached to the previous layers through calibrated spherical spacers installed between the layers. The diameter of the spacers is at least the size of the gap. On the surface of the substrates, longitudinal recesses of a semicircular section are made. A heat reflecting coating with a thickness of 0.1 ... ..0.3 microns with a reflection coefficient of 0.7 ... 0.9 is applied to the outer surface of the microplates and the open areas of the substrate. On the substrates in the places of installation of the spherical spacers, layers of silicon dioxide with a thickness of 0.5 ... 1 μm with a thermal conductivity of 1.4 W / (m · K) are applied. Microplates can be made bimorphic, with the coefficient of thermal expansion (CTE) of the layer facing the substrate less than the CTE of the outer layer. Conductive buses can be made on the surface of the silicon substrate, and microplates can be made electrically conductive.

Brief Description of the Drawings

In FIG. 1 shows the layout of EVI KA microplates on a substrate.

In FIG. 2 shows a cross-section (in the middle of the plates parallel to the horizontal) of the EVI KA with grooves of a rectangular section.

In FIG. 3 shows the cross-section of the EVI KA with grooves of rectangular cross section and longitudinal recesses of a semicircular section.

In FIG. 4 shows a cross section of an EVI spacecraft with trapezoidal grooves.

In FIG. 5 shows the cross section of the EVI KA with trapezoidal grooves and longitudinal recesses of a semicircular section.

In FIG. Figure 6 shows the cross-section of an EVI spacecraft consisting of several layers separated by calibrated spherical spacers.

In FIG. 7 shows a micrograph (x200) of the EVI KA.

The implementation of the invention

The microstructural multilayer EVI KA consists of layers of silicon nitride 1 (Figs. 1–6), grooves with different geometry of walls 2, silicon substrate 3, layers of silicon dioxide 4, layers of heat-reflecting coating 5, vacuum gap 6, microplates 7, longitudinal recesses of semicircular section 8 and calibrated spherical spacers 9.

EVI KA is intended for thermal insulation of the spacecraft surface by protecting it from solar radiation and heat conduction on the spacecraft surface from the surface of the EVI spacecraft. Microplates are rectangular plates with a width of 330-390 microns, a length of 370 ... 420 microns and a thickness of 5 ... 15 microns, made using MEMS technologies of nickel, copper or aluminum with an external heat-reflecting coating with a thickness of 0.1 ... 0.3 microns with a coefficient reflection 0.7 ... 0.9 in the wavelength range of 0.3 ... 3 microns, which accounts for the bulk of the energy of solar radiation. The same heat-reflecting coating is applied to the entire surface of the substrate that is not covered by microplates. Technological openings (not shown in cross sections) with a size of 10 ... 15 μm are made in microplates, which do not significantly affect the thermal properties of the structure.

Rectangular microplates are mounted on a substrate with a gap of 10-20 microns. The gap should be greater than the critical value at which inhomogeneous waves dominate the heat transfer. The critical size of the gap substantially depends on temperature, and at low temperatures this type of heat transfer dominates even at distances of several millimeters. At temperatures from 0 ° C and above, the critical distance varies slightly and is from 8.4 μm or lower, so the value 10 μm is chosen as the lower boundary of the gap. According to studies, a gap of 20 μm prevents heat conduction on the spacecraft surface at temperatures up to –173 ° C [A.I. Volokitin. Radiation heat transfer and “vacuum” friction between nanostructures // Tomsk State University Journal. Itself. state tech. un-that. Ser. Phys.-mat. Nauki, 16 (2002), 129-139], and its increase does not significantly affect the operation of EVI at low temperatures. Based on this, it is impractical to make a gap of more than 20 microns.

If it is necessary to cool the spacecraft due to the bimorph design of micro plates or the electrostatic control system, the gap can be reduced to almost zero. In this case, the micro plates are pressed to the substrate and, if the temperature on the surface of the spacecraft is higher than the temperature of the EVI of the spacecraft, the heat from the surface of the spacecraft flows due to the heat conduction to the EVI of the spacecraft and is emitted into open space.

To reduce radiant heat transfer, the maximum possible outer surface area of the EVI spacecraft (~ 90%) is covered with a heat-reflecting coating with a thickness of 0.1 ... 0.3 μm with a reflection coefficient of up to 0.9. To minimize heat conduction from the EVI spacecraft surface to the spacecraft surface, thermal insulation of the contact points of microplates with a silicon substrate and a non-profiled surface of the reverse side of the substrate is applied by microprofiling the back side of the substrate with grooves of rectangular or trapezoidal section. To further reduce the thermal conductivity of the EVI KA, it is possible to form additional layers of the EVI KA, separated by spherical spacers with low thermal conductivity and point thermal contact with the separated layers.

EVI KA is a silicon substrate 400 ... 500 μm thick with metal microplates made on the surface and additional functional layers using MEMS technologies. The base surface facing the spacecraft is profiled by plasma chemical etching to reduce the area of thermal contact between the microstructured multilayer EVI spacecraft and the outer surface of the spacecraft. It is possible to use both anisotropic deep plasma-chemical etching of silicon (Fig. 2, 3) and an isotropic process of etching silicon (Fig. 4, 5). The etching process affects the entire reverse side of the silicon base, with the exception of areas on which longitudinal recesses of a semicircular section under the spacers are formed. Plasma-chemical etching is carried out through a mask of silicon nitride, formed by photolithography methods, which, after etching, performs the function of thermal insulation of the contact points between the EVI spacecraft and the satellite surface. The thermal conductivity of silicon nitride is 19 W · m / K, which is an order of magnitude lower than that of silicon - 157 W · m / K. Of the standard materials used by the microelectronic industry, silicon nitride is second only to silicon dioxide (1.4 W · m / K) and polymers such as polyimides and PMMA (0.12 and 0.2 W · m / K, respectively) in terms of thermal conductivity. However, silicon nitride has the coefficient of thermal expansion closest to silicon (2.8 10 ^-6 / ° C, for silicon 2.6 10 ^-6 / ° C), which in the conditions of outer space and frequent temperature drops over a wide range is the most critical . Polymers in open space and in the absence of external protection are subject to the destructive effects of atomic oxygen, which significantly reduces the life of the product.

The grooves of rectangular cross section are formed by a depth of 50 ... 250 microns, a width of 40 ... 100 microns, a length of 200 ... 250 microns, the distance between the grooves is 5 ... 10 microns and are located under the fastening points of rectangular microplates on the substrate of the EVI KA layer. The trapezoidal grooves are formed with a depth of 50 ... 250 microns, an outer base width of 100 ... 200 microns, a base width in the substrate body of 90 microns, a length of 200 ... 250 microns, the distance between the grooves of 5 ... 10 microns and are located under the fastening points of rectangular microplates on the substrate.

At the points of contact between the microplates and the silicon substrate, a layer of silicon dioxide is formed with a thickness of 0.5 ... 1 μm with a thermal conductivity of 1.4 W / (m · K) for thermal insulation of the base from microplates.

If necessary, increase the thermal insulation characteristics of the EVI KA form the second and subsequent layers of screen-vacuum insulation, interconnected via calibrated spherical spacers with a diameter of about 40 μm (Fig.6).

The longitudinal recesses of the semicircular section in the silicon base are made for self-organization of spacers between the layers at the stage of assembly of the microstructure multilayer EVI KA. In FIG. 7 shows a photo of a prototype EVI layer.

Microplates can be made bimorphic. Technologically, this is not a big deal. The selection of materials with different values of the coefficients of thermal expansion allows us to calculate the magnitude of the change in the gap depending on the ambient temperature, and, therefore, change the heat-shielding properties of EVI. In this case, the CTE of the lower layer should be less than the CTE of the upper layer.

To change the vacuum gap up to zero, the electrostatic control of the EVI spacecraft can be applied. For this, microplates are made electrically conductive, and conductive busbars and pads are made on the opposite plates of the surface of the silicon substrate. When a potential difference of 0 ... 100 V is applied to the substrate and the micro plates, an electrostatic field arises between them and the micro plates are attracted to the substrate, reducing the gap.

Prototypes of the microstructural multilayer EVI are made; they pass the stage of comprehensive research. Preliminary thermal tests show that, in comparison with the currently used types of EVI, its mass is 10 ... 15 percent less, the thickness is reduced by 20 ... 25 percent.

Claims (7)

1. Microstructural multilayer screen-vacuum insulation of spacecraft, each layer of which is made in the form of a substrate installed on the surface of the apparatus, on which heat-reflecting elements are fixed, characterized in that the heat-reflecting elements are made in the form of an array of rectangular microplates, each rectangular microplate is mounted on a substrate with a gap of 10 ... 20 microns from the surface of the substrate. 2. The microstructural multilayer screen-vacuum insulation of spacecraft according to claim 1, characterized in that the grooves of a rectangular or trapezoidal section are made on the side of the substrate facing the spacecraft. 3. The microstructural multilayer screen-vacuum insulation of spacecraft according to claim 1, characterized in that the second and subsequent layers of screen-vacuum insulation are interconnected via calibrated spherical spacers, while the diameter of the spacers is not less than the gap between the substrate and microplates, and On the side of the substrate facing the spacecraft, longitudinal recesses of a semicircular section are made. 4. The microstructural multilayer screen-vacuum insulation of spacecraft according to claim 1, characterized in that the outer surface of the micro plates and the exposed surfaces of the substrate are coated with a heat-reflecting coating 0.1 ... 0.3 μm thick with a reflection coefficient of 0.7 ... 0.9. 5. The microstructure multilayer screen-vacuum insulation of spacecraft according to claim 3, characterized in that layers of silicon dioxide with a thickness of 0.5 ... 1 μm with a thermal conductivity of 1.4 W / (m · TO). 6. The microstructural multilayer screen-vacuum insulation of spacecraft according to claim 1, characterized in that the microplates are bimorphic, and the coefficient of thermal expansion (KTR) of the layer facing the substrate is less than the KTP of the outer layer. 7. The microstructural multilayer screen-vacuum insulation of spacecraft according to claim 1, characterized in that the surface of the silicon substrate is made of conductive buses, and the micro plates are made of electrically conductive.

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