JPH03253497A - Manufacture of thermal protecting member for space shuttle - Google Patents

Abstract

PURPOSE:To easily obtain a thermal protecting member for a space shuttle by finishing the outer surface of a carbon fiber reinforced carbon complex material into rough surface, and attaching metal silicon powder to the outer surface, and subjecting the complex material to heat treatment so as to generate silicon carbonate on the outer surface to form a covering film of silicon carbonate, and impregnating the film with a mixture of boron oxide and silicon oxide. CONSTITUTION:In manufacturing a thermal protecting member which is a component member of the thermal protecting structure of a space shuttle, the outer surface of a carbon fiber reinforced carbon complex material 4 in which UD sheets comprising carbon fibers aligned in one direction are pseudo-isotropically stacked at an angle of 0 deg./90 deg./+ or -45 deg. is finished into rough surface and then metal silicon powder is attached to the outer surface. The complex material is then subjected to heat treatments under inactive atmosphere and silicon carbonate is initially formed on the outer surface to form a bed layer 5 of silicon carbonate and then a covering layer 8 made from silicon carbonate is formed on the outer surface by vapor-phase chemical reaction deposition method and then the covering layer 8 of silicon carbonate is impregnated with a mixture of boron oxide and silicon oxide so that the thermal protecting member of a space shuttle is manufactured.

Description

DETAILED DESCRIPTION OF THE INVENTION (Field of Industrial Application) The present invention relates to a method for manufacturing a thermal protection member suitable for a spacecraft.

(Prior technology) As a thermal protection system to protect the spacecraft from the high temperatures caused by aerodynamic heating when it re-enters the atmosphere, the U.S. Space Shuttle uses a thermal protection system that protects the spacecraft from the high temperatures caused by aerodynamic heating when it re-enters the atmosphere. Silica tiles are used in all but one area. However, this silica-based tile has low strength and is prone to damage or chipping during use, and its heat resistance is as low as 1280°C, so 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 the outermost layer is made of carbon fiber-reinforced carbon composite material, which is lightweight, has high strength, is resistant to thermal shock, and has excellent heat resistance. However, since the carbon fiber reinforced carbon composite material is entirely composed of carbon, it is easily oxidized and its long-term use in an oxygen-containing atmosphere is limited to 500-600°C.

In order to improve the oxidation resistance of carbon fiber reinforced carbon composite materials,
Some efforts are being made. As an example,
There is a method of impregnating phosphoric acid-based or boron oxide-based glass. This is because the impregnated glass melts during use at high temperatures and covers the external or internal surface of the carbonaceous material to prevent oxidation of the carbonaceous material. In addition, an oxidation-resistant substance (for example, Ti
.

A method has been proposed in which Si, B, W, Ta, and Aj are dispersed in the form of carbides, organic substances, or elements. Furthermore, there is a method of coating the outer surface of the carbon fiber reinforced carbon composite material with a dense film of silicon carbide or silicon nitride obtained by vapor phase chemical reaction deposition method (hereinafter abbreviated as CVD method).

In addition, carbon fiber-reinforced carbon composites can be produced using methods such as the back method in which carbon materials are buried in a mixed powder of alumina, silicon carbide, and metal silicon and heated, and the method in which silicon-containing substances and carbonaceous base materials are directly reacted. A method of producing silicon carbide on the surface of the material has also been proposed.

(Problems to be Solved by the Invention) However, such conventional techniques have the following problems. That is, in the method of impregnating glass with phosphoric acid or boron oxide, when the temperature exceeds about 1000°C, the glass evaporates significantly and cannot become an effective protective film. Even if used in combination with other high-melting-point glasses, phosphoric acid or boron oxide glasses tend to evaporate rapidly at high temperatures, so a long life cannot be expected. In addition, in the method of dispersing oxidation-resistant substances in the matrix, a large amount of oxidation-resistant substances is required to obtain sufficient oxidation resistance, which may cause a decrease in the strength of carbon fiber reinforced carbon composites or the peculiar pseudo-ductility. There are issues such as loss of properties.

In the method of making 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.
The coefficient of thermal expansion of carbon fiber-reinforced carbon composites is about 5X10-'/, whereas the thermal expansion coefficient of carbon fiber-reinforced carbon composites is about 11 to IXI O-'/, and thermal stress causes cracks in the dense film, which leads to Sufficient oxidation resistance could not be obtained due to the intrusion of oxygen. Therefore, attempts were made to seal the cracks with silicon oxide, but since the melting temperature of silicon oxide was as high as 1750°C,
It is not possible to prevent oxygen from entering below the melting temperature of silicon oxide, and satisfactory results have not been obtained. Furthermore, since the film produced by the CVD method is only physically bonded to the base material, it tends to peel off due to thermal shock and lacks reliability. Furthermore, silicon carbide films produced by the bank method or by directly reacting silicon-containing materials with carbon materials lack denseness and cannot be effective oxygen diffusion prevention films.

A carbon composite material in which short carbon fibers are arranged isotropically within the laminated plane has isotropic mechanical properties within the laminated plane, but has a tensile strength of about 10 kgf/ms+'', and is lightweight and It has poor mechanical properties as a material for spacecraft components that require high strength.

(Means for Solving the Problems) Therefore, as a result of intensive studies to solve these problems, the present inventors developed a UD in which carbon fibers are aligned in one direction using a silicon carbide coating film treated with a specific compound. Seat 0°/90
The present inventors have discovered that the above-mentioned problems can be solved by providing the carbon fiber-reinforced carbon composite material on the outer surface of a quasi-isotropically laminated carbon fiber composite material with An object of the present invention is to provide a thermal protection structural member.

This purpose is to provide a thermal protection material whose peripheral edge is formed into a step shape and which is provided on a heat insulating material layer coated on the outer surface of the main body of a space device, and whose stepped portions fit into each other. Thermal protection of a spacecraft, which has a fastening member fixed to the thermal protection member and interposed with a heat insulating layer while the other side is fixed to the space equipment body to fix the heat protection member and the heat insulating layer to the spacecraft body. When manufacturing heat protection members that are structural components, UD sheets with carbon fibers aligned in one direction are heated at 0°/9.
After roughening the outer surface of the carbon fiber-reinforced carbon composite material laminated quasi-isotropically at 0°/±45°, metallic silicon powder is attached and heat-treated in an inert atmosphere to preliminarily coat the outer surface. After producing silicon carbide, forming a coating film made of silicon carbide on the outer surface by a vapor phase chemical reaction deposition method, and then impregnating the silicon carbide film with a mixture of boron oxide and silicon oxide. It is distantly related to the method of manufacturing thermal protection materials for spacecraft.

The thermal protection member obtained by the manufacturing method of the present invention withstands rapid aerodynamic heating during atmospheric reentry and protects the internal insulation layer, so it can also be used to properly protect the spacecraft body wrapped in the insulation layer. It can maintain the temperature.

The method for manufacturing the heat protection member of the present invention will be explained in detail below.

The carbon fiber-reinforced carbon composite material of the present invention has a UD sheet in which carbon fibers are aligned in one direction so that the direction of the carbon fibers is 0° with respect to any direction on the laminated surface of the carbon fiber-reinforced carbon composite material. Direction, 90' direction, +45° direction, -456
It is not particularly limited as long as it is a composite material that is arranged in four directions, that is, pseudo-isotropically laminated, and uses carbon as a matrix (hereinafter abbreviated as carbon fiber reinforced carbon composite material). . For example, carbon fiber (including graphitized fiber)
It is made by molding a UD sheet that is aligned in one direction using a thermosetting resin such as phenolic resin or pitch, and then carbonizing or graphitizing it. It may also be a carbon fiber-reinforced carbon composite material that has been densified by repeating impregnation and carbonization or graphitization with a thermosetting resin or pitch, or by depositing pyrolytic carbon. Further, as the carbon fibers used, fibers generally referred to as carbon fibers, such as polyacrylonitrile carbon fibers, pitch carbon fibers, and rayon carbon fibers, or their precursors are used. Preferably, carbon fiber with a high elastic modulus is used. Further, the plate thickness of the carbon fiber reinforced carbon composite material of the present invention is usually selected from about 0.5 to 100 mm, preferably about 0.7 to 10 ma+.

Next, carbon fiber reinforced carbon composite material (4 in Figure 2)
The surface is roughened. Specifically, a method can be used in which hard particles such as silicon carbide are sprayed onto the surface of the carbon fiber-reinforced carbon composite material using compressed air or the like.

Further, on the surface of the carbon fiber-reinforced carbon composite material, carbon and silicon of the carbon fiber-reinforced carbon composite material are reacted to form a base layer (5) of silicon carbide that is well adhered to the carbon fiber-reinforced carbon composite material. Specifically, a liquid that does not react with metal silicon, for example,
A suspension of metal silicon powder dispersed in isopropyl alcohol is applied to the surface of the carbon fiber reinforced carbon composite material.
The liquid is evaporated to deposit the metallic silicon powder onto the carbon fiber reinforced carbon composite. This is heated in an inert atmosphere to a temperature above the melting point of metal silicon and below 2300''C to cause the carbon of the carbon fiber-reinforced carbon composite material to react with the metal silicon to form a silicon carbide base layer.

The resulting silicon carbide base layer consists of two layers. The particle size of the outer layer is 3-10t! m SiC is a porous layer 20-30 μm thick, slightly integrated at the points of contact between the particles. Under this porous layer (6 in FIG. 3), a mixture layer (7) of silicon carbide and carbon is generated, as if a wedge of silicon carbide was driven into the carbon fiber-reinforced carbon composite material. This is because metal silicon in a molten state penetrates into the pores of the carbon fiber-reinforced carbon composite material that is the base material and reacts with it. The thickness of this mixture layer can be controlled by the amount of metallic silicon deposited before the reaction, and is preferably 550-200 IJ. However, unreacted silicon may remain in the mixture.

A silicon carbide coating film (8 in FIG. 2 or 3) is formed on the silicon carbide underlayer by CVD.

Specific methods include, for example, reducing silicon tetrachloride with hydrogen and reacting with a hydrocarbon such as methane, and thermally decomposing methyltrichlorosilane. C
The thickness of the silicon carbide film formed by the -VD method may be about 10 μm or more, preferably about 100 μm, and usually 50 to 1000 μm.

When silicon carbide is deposited by the CVD method on a silicon carbide base layer, the silicon carbide deposited by the CVD method is also deposited in the pores of the porous silicon carbide layer, so the adhesion of the silicon carbide film by the CVD method to the base material is reduced. Strength improves. The silicon carbide and carbon mixture layer makes this adhesion more reliable. Furthermore, silicon carbide in the mixture layer tends to grow in the pores of the carbon fiber-reinforced carbon composite, and can block the pores near the surface of the carbon fiber-reinforced carbon composite to further reduce the penetration of oxygen into the interior. Be expected. In addition, within 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 CVD method will be alleviated by grading the composition. be done.

The above-described surface roughening treatment, silicon carbide base layer, and silicon carbide coating by CVD method are preferably applied to the entire outer surface of the carbon fiber-reinforced carbon composite material, including the side surfaces.

Finally, the crank formed in the silicon carbide coating film formed by the CVD method is treated with a needle groove treatment using a mixture of boron oxide and silicon oxide (9 in FIG. 2). The melting point of boron oxide is 480°C
is the temperature at which carbon fiber-reinforced carbon composites begin to oxidize (
At temperatures of 500-600'C), boron oxide becomes a liquid and completely penetrates the crank of the silicon carbide film, and at high temperatures where boron oxide significantly evaporates, silicon oxide or borosilicate glass becomes a liquid and cracks. 10) to prevent oxygen from entering through cracks formed in the silicon carbide coating film. The mixture of boron oxide and silicon oxide may be present in the crank of the silicon carbide membrane formed by the CVD method, and there is no problem even if the mixture is present on the silicon carbide membrane or inside the pores of the carbon fiber-reinforced carbon composite material.

The amount of boron oxide is 0.2 to 100 per unit surface area of the carbon fiber reinforced carbon composite material coated with silicon carbide by CVD method.
It is sufficient if it is impregnated with mg/ci, preferably 0.5
~10mg/cil! It is sufficient if it is impregnated.

The amount of silicon oxide may be at least 50%, preferably 1 to 4 times, the weight of boron oxide.

It is also possible to directly impregnate boron oxide or silicon oxide, but
Since the width of cracks in the silicon carbide film produced by the CVD method is narrow, direct impregnation requires high temperature and high pressure equipment, which is not economical. Therefore, a method is suitable in which silicon carbide is impregnated with an organic precursor that has good wettability and is then converted into boron oxide or silicon oxide. One of the organic precursors that satisfies these conditions is a solution of boron or silicon alkoxide, water, and a solvent that can dissolve both.

Specifically, the boron alkoxides include triethyl orthoborate B (0CzHs)sC and below;
It is abbreviated as TEOB. ), and the silicon alkoxide is tetraethyl orthosilicate Si(0CZH5)4
(hereinafter abbreviated as TE01), ethyl alcohol or methyl alcohol can be used as a common solvent. Further, TE01 and TEOB may be subjected to condensation polymerization in advance to such an extent that the viscosity of the solution does not exceed about IF. TEO3/water/ethanol solution or TEO
After the B/water/ethanol solution is impregnated into the object to be treated, it is heat-treated at about 120°C in the atmosphere (hereinafter referred to as hardening treatment).
This results in a compound containing about 80 wt% boron oxide or silicon oxide. There are vacuum impregnation methods in which a container containing carbon fiber reinforced carbon composite material is reduced in pressure, then an organic precursor is introduced under reduced pressure and then returned to normal pressure, and a vacuum pressure impregnation method in which further pressure is applied after vacuum impregnation. A method such as a deimpinging method, in which the object to be treated is simply immersed in an organic precursor solution, can be used.

After completion of the impregnation and hardening treatment with a predetermined organic precursor, heat treatment is performed at 500-1500° C. before use to melt the boron oxide and make the crank needle groove formed by the boron oxide more reliable.

The obtained thermal protection member is constructed by placing a heat insulating material layer (2) made of alumina fiber etc. on the fuselage body (1 in Figure 1) made of aluminum alloy etc. In the airframe protection structure of a spacecraft, the spacecraft is covered with a thin, high-strength heat-resistant material (3) that can support heating and aerodynamic external forces, and this heat-resistant material is fixed to the fuselage body with fastening members such as fasteners. It can be used as In addition to areas where silica tiles have traditionally been used, they are also used in areas that become particularly hot during re-entry into the atmosphere, such as nose cones, leading edges of wings, vertical stabilizers, and body flaps. can do. When used in a nose cone or wing, it is also possible to use it without interposing a heat insulating layer between the heat protection member and the fuselage body.

(Example) Hereinafter, it will be explained in more detail with reference to Examples.

Examples and Comparative Examples A UD sheet phenol prepreg made by aligning carbon fibers in one direction was produced, and this prepreg was layered at +45° layer/9.
0° layer/-45° layer 10° layer 10° layer/-45° layer/9
After laminating 8 layers (0° layer/+45° layer) and forming under pressure and heat,
Firing was performed in a non-oxidizing atmosphere, and then impregnation and firing with a carbon precursor was repeated to obtain a carbon fiber-reinforced carbon composite material with a fiber volume content of 50 vol%. After processing the obtained carbon fiber-reinforced carbon composite material into a predetermined size, silicon carbide powder was sprayed onto the carbon fiber-reinforced carbon composite material using compressed air to roughen the surface of the carbon fiber-reinforced carbon composite material. Next, a suspension of 100 parts of metal silicon powder dispersed in 40 parts of isopropyl alcohol was applied to the surface of the carbon fiber-reinforced carbon composite material, and after evaporating the isopropyl alcohol, it was heated to 2000°C in argon. A base layer of silicon carbide that adhered well to the base carbon fiber-reinforced carbon composite material was created. Subsequently, by CVD method using methyltrichlorosilane,
100μ of SiC was deposited. The above treatment was applied to the entire outer surface of the carbon fiber reinforced carbon composite material.

Next, 3100 parts of TEO and 60 parts of ethanol.

A mixed solution of 26 parts of water and a mixed solution of 1100 parts of TE01, 100 parts of ethanol, and 20 parts of water were alternately impregnated three times each. After impregnation with TEO3 solution or TEOB solution, each was dried and cured at 120°C. The amount of boron oxide impregnated at this time was 1.6 g/cii, and the amount of silicon oxide impregnated was 4.8 g/d. Finally, it was heated to 1000°C under argon.

A test piece of thermal protection material for a spacecraft treated in this way (3
0X30X1.5a+i) in the atmosphere with a heat flux of 0.05
360 kcal/cm”sec argon plasma
The second irradiation test was repeated five times. The weight loss in the example was 0.6 wt%.

A comparative example will be explained below. After impregnating a carbon fiber aggregate cut to a length of 20 mm with phenolic resin, it was heated and pressed from one direction to obtain a shaped rod, then burned out in a non-oxidizing atmosphere, and then impregnated with a carbon precursor. Firing was repeated to obtain a carbon fiber-reinforced carbon composite material with a fiber volume content of 45vol%. Thereafter, test pieces were prepared in the same manner as in the examples.

Table 1 shows the tensile strength at room temperature of Examples and Comparative Examples in which the effective length was 301 and the tab length was 35III.

From the same table, it was found that the examples were superior in mechanical properties. That is, the heat protection member of this example has high strength and the internal carbon fiber-reinforced carbon composite material does not undergo oxidative consumption even upon re-entry into the atmosphere, so it is suitable as a heat protection member for a spacecraft.

Table 1 Comparison of tensile strength

[Brief explanation of drawings]

FIG. 1 is a schematic sectional view showing an example of the use of the thermal protection member of the present invention, FIG. 2 is a schematic sectional view of the thermal protection member for a spacecraft according to the present invention, and FIG. 3 is FIG. FIG. 1: Airframe structural material, 2: Heat insulation layer, 3: Heat protection member, 4:
UD sheet with carbon fibers aligned in one direction at 0°/90
'/±45° pseudo-isotropically laminated carbon fiber-reinforced carbon composite material, 5: silicon carbide base layer, 6: porous layer in silicon carbide base layer, 7: silicon carbide and carbon in silicon carbide base layer Mixture layer, 8: silicon carbide coating layer, 9: mixture of boron dioxide and silicon oxide, 10: molten boron oxide or silicon oxide,
12: Fastener. (Effects of the Invention) According to the present invention, a heat protection member for a spacecraft re-entering the atmosphere can be easily obtained.

Claims (1)

[Claims] (1) A thermal protection member having a stepped peripheral edge formed on a heat insulating material layer coated on the outer surface of the main body of the spacecraft and having stepped portions that fit into each other; A thermal protection structure for a spacecraft having a fastening member fixed to a protective member and interposed with a heat insulating layer while the other side is fixed to the spacecraft body to fix the thermal protection member and the heat insulating layer to the spacecraft body. When manufacturing a thermal protection member that is a component of
After roughening the outer surface of the carbon fiber-reinforced carbon composite material laminated quasi-isotropically at 5°, metallic silicon powder is attached and heat-treated in an inert atmosphere to form silicon carbide on the outer surface in advance. After that, a coating film made of silicon carbide is formed on the outer surface by a vapor phase chemical reaction deposition method, and then the silicon carbide film is impregnated with a mixture of boron oxide and silicon oxide. Manufacturing method for thermal protection materials for shuttle planes.