JPH06135398A - Heat protecting material - Google Patents

Abstract

PURPOSE:To provide a heat protecting material for a space shuttle having a higher adiabatic property compared to conventional heat protecting materials. CONSTITUTION:An adiabatic structure is formed by laminating a plurality of combined sets of heat insulating boards 1-9 and reflecting boards 2-10. A higher-density heat insulating board is applied to the outer heat insulating board 1. A lower-density heat insulating board 9 is applied to the inner beat insulating board. Thus, a lighter and thinner heat protecting material compared to conventional heat protecting materials is available. Blackening the inside surfaces of the reflecting boards 2-10 provides good beat absorption form the inside, further improving adiabatic effect.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a thermal protection material which is installed on the outer surface of a body of a space traffic aircraft and protects the body from aerodynamic heating when the earth reenters the atmosphere.

[0002]

2. Description of the Related Art Space vehicles are exposed to a large amount of heat generated by friction with air when re-entering the atmosphere. Figure 3
Shows the temperature distribution on the surface of the space vehicle during reentry into the atmosphere. The surface of the body excluding the tip of the spacecraft or the tip of the wing is heated up to 1300 ° C. for about 25 minutes.

However, the fuselage itself is made of titanium, aluminum alloy, or CFRP (carbon fiber composite material), and even if they are heat resistant, the maximum operating temperature is 300 ° C. It is necessary to cover the airframe surface of the transfer aircraft with thermal protection material.

The following characteristics are required for the heat protection material for such a space shuttle. (1) Light weight. (2) Be thin. (3) It has excellent heat resistance. (4) Excellent heat insulation properties. (5) Excellent vibration and acoustic resistance. (6) Other excellent manufacturability, economy, reliability, workability, etc.

By the way, various proposals have been made in the past for such a thermal protection material for a space shuttle, but as the prior art closest to the present invention, Japanese Patent Laid-Open No. 1-22 is known.
The disclosure of Japanese Patent No. 02600 can be mentioned.

[0006] In this prior art, as a concept of heat cutoff design at the time of reentry of the atmosphere of a space vehicle, in a high altitude region (65 to 75 km) which is a high heating region, a heat made of a porous heat resistant inorganic material is used. The gas (atmosphere) contained in the protective material suppresses heat transfer and reduces the heat conduction due to radiation from the thermal protective material on the outside, while the low heating area is a low altitude area.
At (55 km or less), the heat stored in the heat protection material in the high altitude region is mainly released to the outside, and as a heat protection material, the outside (heating surface side, atmosphere) The thermal protection material has a laminated structure with multiple layers of heat insulating material so that the density of the heat insulating material increases from the side) to the inside (airframe side).

That is, in this conventional technique, as shown in FIG. 2, a plurality of heat insulating plates 1, 3, 5, 7, 9 and a plurality of reflecting plates 2, 4, 6, 8, 10 are sequentially arranged. The thermal protection material B is configured by a laminated multilayer structure, and at this time, when the densities of the heat insulating plates 1 to 9 are D1 to D9, respectively, the relationship of D9 <D7 <D5 <D3 <D1 is established, As a result, a structure in which a high-density heat insulating material is gradually arranged from the outside to the inside is adopted.

Here, as the heat insulating materials 1 to 9, porous inorganic materials known by trade names such as fine flex, ibiol, and caor are used, and their density is in the form of board or felt. Shape or bulk shape can be arbitrarily selected. A metal plate such as aluminum is used as the reflection plates 2 to 10.

In FIG. 2, reference numeral 11 denotes a carbon / carbon composite material coated with SiC ceramics, which is generally a heat protection material B in consideration of heat resistance and oxidation resistance.
12 is the outer side of the spacecraft, and 12 is the body (cover material) of the space shuttle.

[0010]

In the above prior art, since the high-density portion of the thermal protection material B is inside, the heat generated in the high heating region in the high altitude region when the space vehicle is re-entry into the atmosphere However, no consideration was given to the fact that the heat is mainly stored in the high-density portion of the heat protection material B, and when the low-altitude region, which is the low-heating region, is reached thereafter, the heat stored in the high-density region is reduced. Is discharged to the outside of the fuselage after passing through the low-density heat insulating layer on the outside, while the inside is in direct contact with the fuselage 12, so it also considerably penetrates into the fuselage 12 side. However, there is a problem that the temperature of the airframe structure rises sharply.

As described above, the body structure of the space shuttle is generally made of titanium, aluminum alloy, or heat-resistant CFRP. The maximum operating temperature of these materials is 300 ° C. at most. Therefore, the conventional technique has a big problem due to the temperature rise in the low altitude region.

An object of the present invention is to solve the above-mentioned problems, and to provide a heat protection material for a space shuttle, which can be made lighter and thinner and has excellent heat insulating properties.

[0013]

The above object is to change the density of a plate-like heat insulating member as a heat protection material in the thickness direction, from the maximum density at the outermost side surface portion to the minimum density at the innermost side surface portion. It is achieved by changing.

[0014]

When the space shuttle re-enters the atmosphere of the earth, in the high heating area of high altitude, the heat insulating layer having the maximum density in the outermost surface portion is mainly heated and the heat is mainly stored therein. In this state, when the space vehicle reaches the low heating area of low altitude, the stored heat starts moving to the outside and inside,
The heat transferred to the outside is dissipated to the atmosphere almost immediately, so that a relatively large amount of heat transfer is obtained, and the heat transferred to the inside is transferred through the insulating material containing low density and much air. So much heat transfer does not occur.

Therefore, it is possible to suppress the amount of heat transferred to the fuselage side while ensuring the required heat radiation amount, and to keep the temperature rise of the fuselage at 300 ° C. or less even at the maximum temperature, which makes it lightweight and thin. Moreover, it is possible to obtain a heat protection material having excellent heat insulating properties.

[0016]

The thermal protection material according to the present invention will be described in detail below with reference to the illustrated embodiments. FIG. 1 shows one embodiment of the present invention. In this embodiment, the present invention is carried out by a multi-layer heat insulation system in which a plurality of heat insulation plates are laminated, as in the above-mentioned prior art. It is known that this multi-layer heat insulation method can significantly improve heat insulation when the weight is the same as compared with a single layer heat insulation method.

In FIG. 1, first, A represents a heat protection material according to this embodiment. The heat protection material A is the heat insulating plates 1, 3, 5, 7, as in the prior art described in FIG. Each one of 9 and each of the reflectors 2, 4, 6, 8, and 10 were combined to form one set of heat insulating plates, and a plurality of sets were laminated as shown in the figure. At this time, the heat insulating plate 1 is arranged at the outermost side, and the heat insulating plate 9 is arranged at the innermost side.

Further, in consideration of the fact that, as mentioned above, the outer side (atmosphere side) of the heat protection material A is aerodynamically heated up to 1300 ° C., the oxidation resistance is usually improved by the ceramic coating. The carbon / carbon composite material 11 is arranged, the thermal protection material A is sandwiched between the carbon / carbon composite material 11 and the spacecraft body 12 of the space vehicle, and the heat protection material A is connected by the post 14 via the fasteners 13 and 15 to form the body 1
It is attached to 2.

Here, first, considering the situation at the time of reentry of the space vehicle into the atmosphere, the following can be understood. First, as shown in FIG. 5, the heat load due to aerodynamic heating is 65
It is the largest in a high altitude region of high vacuum degree of up to 75 km, and thereafter gradually decreases, the vacuum degree decreases, and the atmospheric pressure increases, there is almost no heating, and there is a tendency to be rather cooled. Therefore, in order to obtain a thermal protection material that is lightweight, thin, and has excellent heat insulation properties, consider the heat load characteristics during this re-entry, and also fully consider the characteristics of the heat insulation plate, and optimize. It is necessary to plan.

Next, FIG. 4 shows a qualitative tendency of the heat conductivity and the specific heat when the density of the heat insulating plate made of a porous material is changed. Generally, as the density of the heat insulating plate increases, Although the thermal conductivity λ [Kcal / m · h · ° C] decreases, the specific heat C [Kcal / kg ° C] is almost constant. And this tendency is the same regardless of whether the pressure is high or low.

Therefore, in the present invention, as a basic concept of heat insulation, contrary to the prior art shown in FIG. 2, the outer heat insulation plate has a higher density. That is, in the embodiment of FIG. 1, of the heat insulating plates constituting the heat protection material A, the outermost heat insulating plate 1 has the highest density and the innermost heat insulating plate 9 has the lowest density. Is.

Therefore, assuming that the densities of the heat insulating plates 1 to 9 are D1 to D9, respectively, the relationship of D9 <D7 <D5 <D3 <D1 is established in this embodiment as well as in the prior art of FIG. This is because the arrangement of the heat insulating plates 1 to 9 is opposite to that of the conventional technique in this embodiment.

According to the configuration of this embodiment, the following effects can be obtained. That is, the heat that enters when the initial aerodynamic heating load is large at the time of re-entry into the atmosphere is accumulated in the upper heat insulating plate 1 having a large density and therefore a large heat capacity. next,
When it reaches the low altitude region where the aerodynamic heating load decreases, the heating from the carbon / carbon composite material 11 almost disappears, and as a result, the heat accumulated in the heat insulating plate 1 moves to the outside and the inside. At this time, the heat moving toward the inside moves while sequentially heating the low-density heat insulating plates 3 to 9 inside, so that the amount of heat reaching the fuselage 12 can be suppressed and a large temperature can be achieved. The risk of rising can be eliminated.

FIG. 6 shows a thermal protection material A according to the embodiment of FIG.
2 shows the heat insulation analysis result in the completely same state including the dimensions in the heat protection material B according to the related art of FIG.
In the figure, the solid line 16 indicates the carbon / carbon composite material 11.
1, the chain line 17 is the temperature history of the machine body 12 in the embodiment of the present invention in FIG. 1, and the broken line 18 is the temperature history of the machine body 12 in the prior art of FIG. From the result of the heat insulation analysis of FIG. 6, it is understood that the present invention is superior to the conventional technique, and therefore, even when the CFRP is used as the machine body 12, the present invention can sufficiently maintain the safety.

It is to be noted that the heat insulating properties are improved by increasing the density of the heat insulating plates 3 to 9 on the inner side, but an increase in weight cannot be suppressed and desired properties as a heat protection material cannot be obtained. Therefore, in the present invention, the thickness dimension and weight,
In consideration of the balance between the heat insulation characteristics and the heat insulation characteristics, it was found that if a heat insulation plate with a high density is used as the outer heat insulation plate, a sufficient heat insulation function can be obtained even if the inner heat insulation plate has a low density. It adopts a different configuration.

That is, in the embodiment of the present invention, as shown in FIG. 1, the density configuration of the heat insulating plate is set to be higher toward the outer side, so that the heat insulating plate becomes higher in density than the prior art shown in FIG.
For example, when the heat insulating properties are the same, it is possible to obtain the heat protection material A having excellent properties that it is lighter and thinner.

Here, as the heat insulating plates 1 to 9, porous inorganic materials known by trade names such as fine flex, ibiol, and caor may be used. As described above, it is possible to arbitrarily select whether to use a felt shape or a bulk shape, and it is already described that a metal plate such as aluminum may be used as the reflection plates 2 to 10. The present invention can be carried out without being limited to these materials.

FIG. 1 is merely an embodiment of the present invention. Therefore, the number of heat insulating plates and the number of reflecting plates are
Needless to say, it should be implemented by arbitrarily changing it according to the required specifications for the heat protection material. Furthermore, the heat protection material may be composed of one heat insulating plate whose density is changed in the thickness direction. Needless to say.

By the way, in the embodiment of FIG. 1, the inner surfaces of the reflectors 2 to 10 (the body 12 of the space shuttle) are used.
(The surface facing toward the side) may be blackened so that heat transfer in the low altitude region is guided to the outside (carbon / carbon composite material 11 side). According to this embodiment, further weight reduction, The thickness can be reduced.

[0030]

According to the present invention, it is possible to obtain a heat protection material having more excellent heat insulating properties as compared with the prior art. That is, if the temperature difference that must be insulated is the same,
The weight and the thickness can be reduced as compared with the conventional technology.
In addition, when the thickness and weight are the same,
The temperature that must be insulated can be higher.

[Brief description of drawings]

FIG. 1 is a sectional view showing an embodiment of a heat protection material according to the present invention.

FIG. 2 is a cross-sectional view showing an example of a thermal protection material according to the related art.

FIG. 3 is a characteristic diagram showing the surface temperature distribution of the space shuttle when reentering the atmosphere.

FIG. 4 is a characteristic diagram showing a general tendency of thermal conductivity and specific heat depending on the density of the specific heat material.

FIG. 5 is a characteristic diagram showing the tendency of the heat load history due to aerodynamic heating from the start of reentry of the atmosphere of the space vehicle to the landing.

FIG. 6 is a characteristic diagram showing a comparison of thermal insulation characteristic analysis results according to the present invention and a conventional example.

[Explanation of symbols]

A Heat protection material according to one embodiment of the present invention B Heat protection material according to conventional technology 1 High-density heat insulating plate 2 Reflector 3 Medium high-density heat insulating plate 4 Reflector 5 Medium density heat insulating plate 6 Reflector 7 Medium low density Heat insulating plate 8 Reflector 9 Low-density heat insulating plate 10 Reflector 11 Carbon / carbon composite material 12 Space vehicle body 13 Upper fastener 14 Post 15 Lower fastener 16 Temperature history of carbon / carbon composite material 17 Space transportation by the present invention 17 Temperature history of the fuselage 12 of the aircraft 18 Temperature history of the fuselage 12 of the space shuttle according to the related art.

 ─────────────────────────────────────────────────── ─── Continuation of the front page (72) Inventor Keiho Kojima 4026 Kujimachi, Hitachi City, Ibaraki Prefecture Hitate Manufacturing Co., Ltd.Hitachi Laboratory Ltd. In Hitachi Research Laboratory (72) Inventor Takatoshi Yoshioka 4026 Kuji Town, Hitachi City, Ibaraki Prefecture Hitate Manufacturing Co., Ltd. (72) Inventor Takehiko Yoshida 4026 Kuji Town, Hitachi City, Ibaraki Hitachi Research Co., Ltd. In-house (72) Toshio Hattori, 502 Jinritsu-cho, Tsuchiura-shi, Ibaraki Machinery Research Laboratory, Hiritsu Manufacturing Co., Ltd. (72) In-house Masanobu Kirihara 3-1-1, Saicho-cho, Hitachi, Hitachi Ibaraki Hitachi Factory (72) Inventor Miki Morino 1 1-1-2, Sengen, Tsukuba-shi, Ibaraki Ushiro Development Corporation Tsukuba Space Center (72) Inventor Tomoyuki Kobayashi 2-4-1, Hamamatsucho, Minato-ku, Tokyo Within the space development business group (72) Inventor Toshinari Yoshinaka 1 Uchiku Development Corporation Tsukuba, 1-chome, 2-chome, Sengen, Tsukuba, Ibaraki Inside the Space Center

Claims (4)

[Claims] 1. A thermal protection material for a space shuttle, which comprises a porous plate-like heat insulating member, wherein the density of the plate-like heat insulating member is changed in the thickness direction so that the maximum density at the outermost side surface portion changes from the maximum density at the innermost side surface portion. A thermal protection material characterized by being configured so as to change toward the minimum density of. 2. The heat protection material according to claim 1, wherein the plate-shaped heat insulating member is composed of a laminate of a plurality of heat insulating plates having different densities. 3. The heat protection material according to claim 2, wherein a reflection plate is provided between a plurality of heat insulating plates that form the laminate. 4. The heat protection material according to claim 3, wherein the inner surface of the reflection plate is processed to have a black color or a color close to black.