RU173721U1 - Scheme of a heat-shielding coating of a reusable heat shield of a descent vehicle for returning from a low Earth orbit - Google Patents

Abstract

The utility model relates to reusable heat-shielding composite coatings that can be used in space technology. The technical result of the proposed utility model is a heat-shielding coating scheme for the reusable heat shield of the descent spacecraft, optimized for returning the device from low Earth orbit (for example, from the International Space Station (ISS)). The technical result is achieved in that the heat-shielding coating scheme of the reusable heat shield of the descent spacecraft for returning from a low Earth orbit is characterized by a composition of layers of a non-woven carbon skeleton and a silicon carbide matrix and relates to coatings of substantially anisotropic materials with different physical, mechanical and thermal characteristics as layers and the thickness of the material. Moreover, for returning the apparatus from a low Earth orbit, the coating has three layers: an outer dense layer with a residual porosity of less than 5%, a layer of carbon-ceramic composite material (UKKM) 2 mm thick, a second layer of UKKM 50% porosity 15 mm thick and a third layer from a high-temperature fiber heat insulator brand Saffil (manufactured by DuPont) with a thickness of 45 mm. 1 s.p. f-ly, 4 ill.

Description

Technical field

The utility model relates to reusable heat-shielding composite coatings that can be used in space technology.

State of the art

Known RF patent No. 164403 (IPC F16L 59/02, B32B 18/00, published 08/27/2016) for the author's utility model "SCHEME OF HEAT PROTECTIVE COATING BASED ON GRADIENT POROUS CARBON-CERAMIC COMPOSITE MATERIAL" (adopted as the closest). This heat-protective coating scheme (TZP) is characterized by a composition of woven and non-woven carbon skeleton layers and a matrix containing carbon and silicon carbide components, and refers to coatings of substantially anisotropic materials with different physical, mechanical and thermal characteristics of both layers and thickness material.

However, this coating scheme is unnecessarily complicated, as it consists of five layers arranged in series: the first outer (front) layer of carbon-ceramic composite material (UKKM) based on a woven carbon frame with a residual porosity of less than 5% and a thickness of 3 to 5 mm to ensure thermal and oxidative stability, the second layer under the first layer of at least 5 mm thick of a carbon-ceramic composite material based on a non-woven carbon frame with a residual porosity of at least e 18%, the third layer under the second layer of at least 5 mm thick of a carbon-ceramic composite material based on a non-woven carbon frame with a residual porosity of at least 22%, the fourth layer under the third layer of at least 5 mm thick of a carbon-ceramic composite material based on a non-woven carbon frame with a residual porosity of at least 38%, the fifth layer under the fourth layer with a thickness of at least 5 mm from a carbon-ceramic composite material based on a non-woven carbon frame with a residual porosity of e less than 48%. Such a complex coating structure leads to a significant increase in the duration and cost of manufacture.

Utility Model Disclosure

The technical result of the proposed utility model is a heat-shielding coating scheme for the reusable heat shield of the descent spacecraft, optimized for returning the device from low Earth orbit (for example, from the International Space Station (ISS)).

The technical result is achieved in that the heat-shielding coating scheme of the reusable heat shield of the descent spacecraft for returning from a low Earth orbit is characterized by a composition of layers of a non-woven carbon skeleton and a silicon carbide matrix and relates to coatings of substantially anisotropic materials with different physical, mechanical and thermal characteristics as layers and the thickness of the material. Moreover, for returning the apparatus from a low Earth orbit, the coating has three layers: an outer dense with a residual porosity of less than 5% UKKM layer 2 mm thick (in the prototype: an outer (front) layer of UKKM based on a woven carbon frame with a residual porosity of less than 5% and thickness from 3 to 5 mm to ensure heat and oxidation resistance), the second layer of UKKM 50% porosity of a thickness of 15 mm (in the prototype: the fifth layer with a thickness of at least 5 mm from UKKM based on a non-woven carbon frame with a residual porosity of at least 48 %) third fibrous layer from a high grade insulator Saffil (manufactured by DuPont, USA) 45 mm thick.

As in the prototype, carbon fibers from artificial cellulose fibers, carbon fibers from PAN fibers and carbon fibers from pitch can be used as a filler for a heat-protective material based on a gradient porous UKCM.

List of figures

FIG. 1 is a geometric model of an element of a representative volume of UKKM used to calculate the thermal conductivity of UKKM taking into account radiation heat transfer in the material;

FIG. 2 - structure of TZP;

FIG. 3 - the density of the heat flux acting on the TZZ during the descent of the vehicle to the Earth from the low Earth orbit of the ISS;

FIG. 4 - change in the maximum (on the outer surface of the TZP) and the minimum temperature (on the inner surface of the TZP lying on the surface of the power structure of the spacecraft) from time to time when the spacecraft descends from the low Earth orbit of the ISS.

Utility Model Implementation

A feature of the porous UKKM is that two processes of heat transfer simultaneously and interconnected in the bulk of the material proceed - conductive along the solid frame and radiation in the space between the fibers. In this case, the ratio of the conductive and radiative heat fluxes depends on the porosity of the material and temperature. For porous media, the total heat transfer can be characterized by an effective coefficient of thermal conductivity, which is the ratio of the density of the total conductive and radiation heat flux passing through the porous medium to the local temperature gradient. To carry out this calculation, a geometric model of a representative volume element with a size of 100 × 50 × 50 μm was chosen (Fig. 1). A porous UKKM model was placed between the plates. The thermal contact of the plates with the fibers of the porous material was considered ideal. A representative element of the volume was surrounded by a shell of zero thickness, which played the role of a mirror, its reflection coefficient was set to 1. The surfaces of the fibers and outer linings were taken gray, diffusely reflecting and their degree of blackness was adopted 0.8. Further, on the basis of the geometric model thus refined, a finite element model was constructed for computer simulation of temperature fields by the thickness of the representative volume element. As a result of modeling the heat transfer process, the value of the total heat flux passing through both the porous material and through the plates was determined. The thermal conductivity of the representative element of the volume of the system "representative element of the volume-lining" is calculated as the sum of the thermal resistances of a portion of a thermal circuit of constant cross section:

where l, l ₁ , l ₂ , l ₃ are the total thickness and thickness of individual sections of the geometric model, respectively, l ₁ = l ₃ = l ^* , and λ, λ ₁ , λ ₂ , λ ₃ are the thermal conductivity of the entire model and the thermal conductivity its individual parts, respectively, λ ₁ = λ ₃ = λ ^* . Then the thermal conductivity of the representative element of volume λ _{2 is} calculated as

The option of thermal loads on the heat-transfer zone was used during the descent from the orbit of the ISS [taken from Walker, SP Preliminary Development of a Multifunctional Hot Structure Heat Shield / SP Walker, K. Daryabeigi, JA Samareh, SC Armand, SV Perino // AIAA Paper 2014-0350 . - 2014. - 13 p], presented in FIG. 3. The descent from the orbit of the ISS lasts about 800 s, the maximum value of the heat flux density is about 110 W / cm ² . In the simulation, it was believed that after the descent was completed, the thermal current transformer cools down for 5000 s.

For thermal design, a multilayer TZP model was used. The thermal load in the descent section was brought to the frontal surface. The radiation heat removal from the front surface of the heat-condensing element into the surrounding space was taken into account.

The following TZP design was simulated. The outer layer of TZP was a dense UKKM layer with a residual porosity (P) of less than 5% 2 mm thick, the role of which is to provide oxidative and erosive protection of the inner porous layers. Behind it was a layer of UKKM. To create the proposed TZP, UKKM with a porosity of 50% was selected. A layer of high-temperature heat-insulating material of the Saffil brand (manufactured by DuPont, USA), which is located on the power structure of the descent vehicle (CA), is placed behind the porous UKKM layer. The power structure of the SA in computer simulation was a layer of aluminum alloy D16 with a thickness of 2.5 mm. In FIG. 2 numbers are marked: 1 - dense UKKM with residual porosity of less than 5%, 2 - porous UKKM (P = 50%), 4 - fibrous heat insulator type Saffil, 5 - aluminum alloy D16.

The purpose of thermal design was to determine the thicknesses of the layers of the porous UKKM and Saffil material, which protect the power structure from overheating. The maximum permissible design temperature of the aluminum alloy D16 was set equal to 165 ° C. UKKM temperature should not exceed 2000 ° C. The criterion for optimization was the minimum linear mass of the heat-transfer unit.

An analysis of the results showed that when designing the TZP SA for returning from the ISS, there is no need to use UKKM layers with variable porosity. The best linear mass results are obtained using UKKM with a porosity of 50%. This result is explained by the fact that the maximum temperature of the porous UKKM is in the range of permissible temperatures of 1700-1850 ° C (see the graphs of Fig. 4: the upper graph is the maximum temperature of the outer layer of the thermal protection layer, the lower graph is the minimum temperature on the surface of the power structure CA).

The table shows the parameters of the optimal variant of the TZP scheme for launching the spacecraft from the ISS low Earth orbit.

The proposed utility model was created and tested during the work under the agreement on the provision of subsidies 14.577.21.0099 between the Ministry of Education and Science of the Russian Federation and MSTU. N.E. Bauman.

Claims (2)

1. The heat-shielding coating of a reusable heat shield of a descent spacecraft, characterized by a composition of layers of a non-woven carbon skeleton and a silicon carbide matrix and relating to coatings of substantially anisotropic materials with different physical, mechanical and thermal characteristics of both layers and material thickness, characterized in that that for the apparatus returned from a low Earth orbit, the coating has three layers: an outer dense carbon-ceramic layer with a residual porosity of less than 5% Cesky composite material (UKKM) with a thickness of 2 mm, a second layer of UKKM with a porosity of 50% and thickness of 15 mm and a third layer of high-grade fibrous heat insulator Saffil manufactured by DuPont 45 mm. 2. The coating according to claim 1, characterized in that carbon fibers from artificial cellulose fiber, carbon fibers from PAN fiber and carbon fibers from pitch are used as filler of the porous UKKM.

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