JP2651386B2 - Thermal protection structure for space equipment - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Industrial Application Field] The present invention relates to a thermal protection structure for space equipment particularly suitable for a spacecraft.

[Conventional technology]

As a thermal protection system to protect the airframe from the high temperature caused by aerodynamic heating when the spacecraft re-enters the atmosphere, the U.S. Space Shuttle uses silica, except for particularly hot parts such as the nose cap and wing leading edge. System tiles are used. However, these silica tiles have low strength, causing problems such as damage and chipping during use. The heat-resistant temperature is as low as 1280 ° C, and the development of a high-strength thermal protection system that can be used at higher temperatures is awaited. Therefore, a thermal protection system has been proposed in which a carbon fiber reinforced carbon composite material that is lightweight, high-strength, resistant to thermal shock and excellent in heat resistance is disposed on the outermost layer. However, since all carbon fiber reinforced carbon composites are composed of carbon, they are easily oxidized and long-term use in an oxygen-containing atmosphere is limited to 500 to 600 ° C.

Some efforts have been made to improve the oxidation resistance of carbon fiber reinforced carbon composites. One example is a method of impregnating with phosphoric acid or boron oxide glass. This is because the impregnated glass melts during use at high temperatures and covers the outer or inner surface of the carbonaceous material to prevent strengthening of the carbon material. In addition, oxidation-resistant substances (for example, T
A method of dispersing (i, Si, B, W, Ta, Al) in a state of carbide, organic matter or element has been proposed. Furthermore, there is a method of coating the outer surface of a carbon fiber reinforced carbon composite with a dense silicon carbide or silicon nitride film obtained by a gas phase chemical reaction deposition method (hereinafter abbreviated as a CVD method). In addition, a carbon fiber reinforced carbon composite material is prepared by a packing method in which a carbon material is buried in a mixed powder of alumina, silicon carbide, and metal silicon and heated, or a method in which a silicon-containing material is directly reacted with a carbonaceous substrate. For example, a method of generating silicon carbide on the surface of silicon has been proposed.

[Problems to be solved by the invention]

However, such a conventional technique has the following problems. In other words, in the method of impregnating glass of phosphoric acid or boron oxide, when the temperature exceeds about 1000 ° C., the evaporation of the glass cannot be a very effective protective film. Even if it is used in combination with another glass having a high melting point, the phosphoric acid or boron oxide-based glass is highly evaporated at a high temperature and a long life cannot be expected. Also, in the method of dispersing the oxidation-resistant substance in the matrix, a large amount of the oxidation-resistant substance is required to obtain sufficient oxidation resistance. There are problems such as loss of properties.

In the method of forming a dense silicon carbide or silicon nitride coating film by the CVD method, the thermal expansion coefficient of silicon carbide or silicon nitride is 3.5
Whereas a × 10 ^-6 / ° about K, the thermal expansion coefficient of the carbon fiber-reinforced carbon composite material is -1 to 1 × 10 ^-6 / ° K, a crack occurs in dense film by thermal stress However, sufficient oxygen resistance cannot be obtained because oxygen enters from here. Attempts were made to seal the cracks with silicon oxide. However, since the melting temperature of silicon oxide was as high as 1750 ° C, sufficient results could not be obtained because oxygen intrusion could not be prevented below the melting temperature of silicon oxide. . Furthermore, since the film formed by the CVD method is only physically bonded to the substrate, it is easily peeled off due to thermal shock or the like, and lacks reliability. Further, a silicon carbide film formed by a pack method or by directly reacting a silicon-containing substance with a carbon material lacks denseness and cannot be an effective oxygen diffusion preventing film.

[Means for solving the problem]

The present inventors have conducted intensive studies to solve these problems, and as a result, by providing a silicon carbide coating film of a specific shape treated with a specific compound on the outer surface of the carbon fiber reinforced carbon composite material, The inventors have found that the problem can be solved, and have reached the present invention. That is, an object of the present invention is to provide a thermal protection structure of a space device suitable for a space device such as a space shuttle vehicle to re-enter the atmosphere.

The purpose is to provide a heat insulating material layer applied on the outer surface of the main body of the space device and a heat protection layer provided on the heat insulating material layer so that the peripheral edge is formed in a stepped shape and the steps are engaged with each other. A silicon carbide coating film is formed on an outer surface of a carbon fiber reinforced carbon composite material whose surface is roughened, and carbon fiber reinforced between the silicon carbide coating film and the carbon fiber reinforced carbon composite material. A heat protection member having a silicon carbide layer obtained by reacting with carbon of the composite material and further having a silicon carbide coating film sealed with a mixture of boron oxide and silicon oxide; and one side fixed to the heat protection member This is achieved by providing a space device heat protection structure having a heat protection member and a fastening member fixing the heat insulation material layer to the space device body while the other side is fixed to the space device body while interposing the heat insulation material layer. .

Hereinafter, the present invention will be described. The present invention includes a heat insulating material layer on a space device main body, and a heat protection member fixed to the space device main body by a fastening member outside the heat insulating material layer. This thermal protection member withstands rapid air heating upon re-entry into the atmosphere and protects the internal heat insulating material layer, so that the space equipment body wrapped by the heat insulating material layer can be kept at an appropriate temperature.

 Hereinafter, the heat protection member will be described in detail.

The carbon fiber reinforced carbon composite material in the present invention is not particularly limited as long as it is a composite material using carbon fiber as a reinforcing material and carbon as a matrix. For example, carbon fibers (including graphitized fibers) are formed using a thermosetting resin such as a phenol resin or pitch, and carbonized or graphitized. Alternatively, a carbon fiber reinforced carbon composite material obtained by repeating impregnation and carbonization or graphitization with a thermosetting resin or pitch or densifying by depositing pyrolytic carbon may be used. As the carbon fiber to be used, a fiber generally called a carbon fiber such as polyacrylonitrile-based carbon fiber, pitch-based carbon fiber, rayon-based carbon fiber, or a precursor thereof is used. The reinforcing form of the carbon fiber is not particularly limited, and may be any form such as lamination of cloth, three-dimensional woven fabric, or short fiber form.

In the present invention, first, the surface of the carbon fiber reinforced carbon composite material (4 in FIG. 2) is subjected to unevenness treatment. Specifically, a method of spraying hard particles such as silicon carbide on the surface of the carbon fiber reinforced carbon composite material with compressed air or the like can be used.

Next, carbon and silicon of the carbon fiber reinforced carbon composite are reacted on the surface of the carbon fiber reinforced carbon composite to form an underlayer (5) of silicon carbide which is well bonded to the carbon fiber reinforced carbon composite. Specifically, a suspension in which metal silicon powder is dispersed in a liquid that does not react with metal silicon, for example, isopropyl alcohol, is applied to the surface of the carbon fiber reinforced carbon composite material, and the liquid is evaporated to form a metal silicon powder. To the carbon fiber reinforced carbon composite. 1700 ~ in an inert atmosphere
Heat to 2300 ° C to react the carbon of the carbon fiber reinforced carbon composite with metallic silicon to form an underlayer of silicon carbide.

The resulting silicon carbide underlayer consists of two layers. The outer layer is a porous layer having a thickness of 20 to 30 μm, in which SiC having a particle system of 3 to 10 μm has a slightly integrated contact point between particles. Under the porous layer (6 in FIG. 3), a mixture layer of silicon carbide and carbon (7) is formed as if a wedge of silicon carbide was driven into a carbon fiber reinforced carbon composite material. This is because the metallic silicon in the molten state penetrates into the pores of the carbon fiber reinforced carbon composite material as the base material and reacts. The thickness of this mixture layer can be controlled by the amount of metallic silicon deposited before the reaction, and preferably 100 to 200 μm. However, unreacted silicon may remain in the mixture layer.

A silicon carbide coating film (8 in FIG. 2 or 3) is formed on the silicon carbide underlayer by a CVD method. Specifically, for example, a method of reducing silicon tetrachloride with hydrogen to react with a hydrocarbon such as methane, a method of thermally decomposing methyltrichlorosilane, and the like can be used. The thickness of the silicon carbide film formed by the CVD method may be about 10 μm or more, preferably about 100 μm, and usually 50 to 1000 μm.

When silicon carbide is deposited by CVD on the silicon carbide underlayer, the silicon carbide deposited by CVD also deposits in the pores of the porous silicon carbide layer. Power improves. A layer of a mixture of silicon carbide and carbon makes this adhesion more secure. Further, the silicon carbide of the mixture layer is easily generated in the pores of the carbon fiber reinforced carbon composite material, and is expected to close the pores near the surface of the carbon fiber reinforced carbon composite material and further reduce the penetration of oxygen into the interior. Is done. In the mixture layer, the ratio of silicon carbide to carbon decreases toward the inside of the base material, so it is expected that the thermal stress generated in the silicon carbide coating film by the CVD method will be reduced by the composition gradient. Is done.

It is preferable that the above-mentioned unevenness treatment, the silicon carbide underlayer, and the silicon carbide coating by the CVD method be applied to the entire outer surface including the side surfaces of the carbon fiber reinforced carbon composite material.

Finally, the cracks generated in the silicon carbide coating film by the CVD method are subjected to groove sealing with a mixture of boron oxide and silicon oxide (9 in FIG. 2). The melting point of boron oxide is 480 ° C, and the temperature at which carbon fiber reinforced carbon composite material starts to oxidize (500
(Up to 600 ° C), the boron oxide becomes liquid and completely closes the cracks in the silicon carbide film. At high temperatures where the boron oxide evaporates significantly, the silicon oxide or borosilicate glass becomes liquid and completes the cracks. (10) to prevent oxygen from entering through cracks formed in the silicon carbide coating film. The mixture of boron oxide and silicon oxide only needs to be present in the cracks of the silicon carbide film formed by the CVD method, and there is no problem if it is present on the silicon carbide film or in the pores of the carbon fiber reinforced carbon composite material.

Boron oxide is per unit surface area of the carbon fiber reinforced carbon composite material coated with silicon carbide by the CVD method, only to be impregnated 1 to 200 mg / cm ^2, if preferably is impregnated 0.10~100mg / cm ² Good. Silicon oxide is 10% of boron oxide by weight
Above, preferably 30 to 400%.

Boron oxide or silicon oxide may be directly impregnated, but since the width of cracks in the silicon carbide film formed by the CVD method is narrow, direct impregnation requires high-temperature and high-pressure equipment and is not economical. Therefore, a method of impregnating a low-viscosity silicon carbide and an organic precursor having good wettability and then converting the same into boron oxide or silicon oxide is suitable. One of the organic precursors that satisfies such conditions is a solution of boron or silicon alkoxide, water, and a solvent capable of dissolving both.

Specifically, triethyl orthoborate B (OC ₂ H ₅ ) ₃ (hereinafter abbreviated as TEOB) is used as the boron alkoxide, and tetraethyl orthosilicate Si (OC ₂ H ₅ ) is used as the silicon alkoxide. ₄ ) (hereinafter abbreviated as TEOS), and ethyl alcohol or methyl alcohol can be used as the common solvent. Also, T
EOS or TEOB may be prepolymerized so that the viscosity of the solution does not exceed about 1P. The TEOS / water / ethanol solution or TEOB / water / ethanol solution is impregnated with the object to be treated and then heat-treated at about 120 ° C. in the air (hereinafter referred to as “hardening treatment”) to obtain an oxide of about 80 wt%. It becomes a compound containing silicon or silicon oxide. The container containing the carbon fiber reinforced carbon composite material is evacuated, followed by a vacuum impregnation method in which the organic precursor is introduced under reduced pressure and then returned to normal pressure, a vacuum impregnation method in which further pressure is applied after vacuum impregnation, A dipping impregnation method in which an object to be treated is simply immersed in an organic precursor solution can be used.

After the impregnation and curing treatment of the predetermined organic precursor is completed, heat treatment is performed at 500 to 1400 ° C. before use to melt the boron oxide, thereby making the cracks of the boron oxide more reliable.

The heat protection member obtained is obtained by disposing a heat insulating material layer (2) made of alumina fiber or the like on an airframe structural material (1 in FIG. 1) made of, for example, an aluminum alloy or a polyimide-based composite material. Is covered with a thin high-strength heat-resistant material (3) that can support aerodynamic external force, and this heat-resistant material is fixed to the fuselage structural material with fasteners or other fasteners, which can be used as a heat protection structure for space shuttle vehicles. . In addition, it is used not only in areas where conventional silica tiles were used, but also in areas that are particularly hot when re-entering the atmosphere, such as nose cones, wing leading edges, vertical tails, body flaps, etc. be able to. When used for a nose cone or a wing, it can be used without a heat insulating layer between the thermal protection member and the body of the machine.

〔Example〕

 Hereinafter, the present invention will be described in more detail with reference to examples.

FIG. 1 shows a thermal protection structure for a spacecraft as one embodiment of the present invention. A heat insulating material layer 2 made of a mixed fiber of silica fiber and alumina fiber is provided on the surface of the main body 1 of the spacecraft, and the outer surface of the heat insulating material layer 2 is covered with a heat protection member 3 to be described later. Although the heat protection member 3 has a tile shape, the end portion of the heat protection member 3 is formed in a step shape, and the step portion overlaps with the adjacent heat protection member 3.
A niobium fastener 12 as a fastening member is inserted into the central portion and one end of the thermal protection member 3, respectively, and the thermal protection member 3 is fixed to the main body 1 of the spacecraft while interposing the heat insulating material layer 2. I have. The lower end of the fastener 12 is fixed by a clip 13 provided on the main body 1 of the spacecraft. The thermal protection member 3 used in this example was manufactured by the method described below.

First, silicon carbide powder was sprayed with compressed air on a flat carbon fiber reinforced carbon composite material laminated with carbon fiber cloth,
The surface of the carbon fiber reinforced carbon composite was made uneven. Subsequently, a suspension obtained by dispersing 100 parts of metal silicon powder in 40 parts of isopropyl alcohol was applied to the surface of the carbon fiber reinforced carbon composite material, and after the isopropyl alcohol was evaporated,
Heating to 2,000 ° C in argon produced an underlayer of silicon carbide that adhered well to the base carbon fiber reinforced carbon composite. Next, the silicon was deposited by CVD using methyltrichlorosilane.
C was deposited 100 μm. The above process was performed on the entire outer surface of the carbon fiber reinforced carbon composite material. Next, a mixed solution of 100 parts of TEOS, 60 parts of ethanol and 26 parts of water, and 100 parts of TEOB and ethanol
A mixed solution of 100 parts and 20 parts of water was alternately impregnated three times each. After impregnation with TEOS solution or TEOB solution, each was dried and cured at 120 ° C. The boron oxide impregnation amount at this time was 16 mg / cm ² , and the silicon oxide impregnation amount was 48 mg / cm ² . Finally, it was heated to 800 ° C. in argon.

As a comparative example, a thermal protection prepared in the same manner as in the example except that a silicon carbide coating film was directly formed on a carbon fiber reinforced carbon composite material by a CVD method without forming a silicon carbide underlayer without performing unevenness treatment. Members were also made.

A test piece (30 × 30 × 5 mm) of the thermal protection member for spacecraft was treated in this way as an argon plasma with a heat flux of 0.05 Kcal / cm ² sec in air as a condition to simulate aerodynamic heating when re-entering the atmosphere. Was repeated 10 times. In the example, the weight loss was 0.27 wt%, whereas in the comparative example, the silicon carbide coating film was peeled off by one argon plasma irradiation.

Therefore, in the case of the thermal protection structure to which the thermal protection member of the present embodiment is attached, even when re-entering the atmosphere, the situation in which the internal carbon fiber reinforced carbon composite material is oxidized and lost due to peeling of the silicon carbide coating film does not occur. Therefore, sufficient heat insulation can be secured, and the spacecraft can be kept at room temperature.

〔The invention's effect〕

ADVANTAGE OF THE INVENTION According to this invention, the thermal protection structure of space equipment suitable as a thermal protection structure for space shuttle vehicles reentering the atmosphere can be obtained.

[Brief description of the drawings]

FIG. 1 is a schematic sectional view showing an example of use of the thermal protection of the present invention, FIG. 2 is a schematic sectional view of the thermal protection for a spacecraft in the present invention, and FIG. It is an enlarged view. 1 ... Airframe structural material, 2 ... Thermal insulation layer, 3 ... Thermal protection member, 4 ...
... Carbon fiber reinforced carbon composite material, 5 ... SiC underlayer, 6 ...
Porous layer in silicon carbide underlayer, 7 ... mixture layer of silicon carbide and carbon in silicon carbide underlayer, 8 ... silicon carbide coating film, 9 ...
A mixture of boron oxide and silicon oxide, 10 molten boron oxide or silicon oxide, 12 fasteners.

 ──────────────────────────────────────────────────続 き Continuing on the front page (72) Inventor Masayuki Oshima 10 Oecho, Minato-ku, Nagoya, Aichi Prefecture Mitsubishi Heavy Industries, Ltd. Address: Mitsubishi Heavy Industries, Ltd., Nagoya Aircraft Works (72) Inventor: Osamu Fujishima 1st Bancho, Sakaide-shi, Kagawa Prefecture Mitsubishi Kasei Co., Ltd. Within the Research Institute, Inc. (72) Inventor Tasuke Nose 1st Bancho, Sakaide-shi, Kagawa Prefecture Mitsubishi Sakaide Plant

Claims (1)

(57) [Claims] A heat insulating material layer provided on an outer surface of a main body of a space device, and a heat insulating material layer provided on the heat insulating material layer, the peripheral edge portion of which is formed in a stepped shape and the steps are fitted to each other. A silicon carbide coating film is formed on an outer surface of a carbon fiber reinforced carbon composite material whose surface is roughened, and a carbon fiber is provided between the silicon carbide coating film and the carbon fiber reinforced carbon composite material. A heat protection member having a silicon carbide layer obtained by reacting with carbon of the reinforced composite material and further having a silicon carbide coating film sealed with a mixture of boron oxide and silicon oxide; A heat protection structure for a space device, comprising: a heat protection member fixed to the space device body on the other side while fixing the heat insulation material layer and a fastening member fixing the heat insulation material layer to the space device body.