JPH06255598A - Heat control panel for space use - Google Patents

Abstract

PURPOSE:To provide a heat control panel for space reducing the temperature difference of a spacecraft by a large margin while being of simple constitution. CONSTITUTION:A heat control panel 1 is fitted to a position of covering the high temperature part 11 and low temperature part 10 of a spacecraft 2, intersecting the sunlight so as to intercept the sunlight while having large sunlight absorptivity on the spacecraft side face and small emissivity on the same face.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a space thermal control panel applied to passive thermal control of a spacecraft.

[0002]

BACKGROUND OF THE INVENTION Heretofore, there has been no space thermal control panel that reflects cosmic heat radiation. Therefore, as shown in FIG. 9, the spacecraft 2 flying on the earth 6 in a predetermined orbit 7 has a portion where the sun rays 3 hit well, a high temperature portion 11, a portion where they do not hit much, and a low temperature portion 10. In the figure, 4 indicates earth emission, 5 indicates earth's solar reflection, and albeto.

[0003]

In the conventional spacecraft 2 described above, the high temperature portion 11 and the low temperature portion 10 depend on the conditions of the orbit 7.
Exists and a temperature difference occurs.

Further, comparing the daytime on the orbit with the sunlight 3 and the nighttime with the shadow of the earth 6 without the sunlight 3, the temperatures are high during the daytime and low during the nighttime. Even in a portion, a temperature difference occurs with time.

Due to these temperature differences, problems such as the generation of thermal stress have occurred.

[0006]

The present invention takes the following means in order to solve the above problems.

That is, as a heat control panel for space,
It should be installed at a position that intersects with the sun's rays and covers the high temperature part and the low temperature part of the spacecraft, shields the sun's rays, and has a large solar ray absorption rate on the spacecraft side and a small emissivity on the same side. did.

[0008]

By the above means, the sun rays hit the heat control panel. Therefore, the incidence on the high temperature part and the low temperature part of the spacecraft, which becomes high and low without the heat control panel, is blocked.

On the other hand, sunlight reflected light from the earth side hits the surface of the heat control panel having a large absorptance and is mostly absorbed. In addition, the earth radiation from the earth hits the thermal control panel, which has a low emissivity, that is, a high reflectance to the earth radiation, and is mostly reflected. This reflected earth radiation hits the hot and cold parts of nearby spacecraft and heats them.

Therefore, each part of the spacecraft is always maintained at a substantially uniform temperature, and the occurrence of defects due to the temperature difference is prevented.

[0011]

【Example】

 (1) A first embodiment of the present invention will be described with reference to FIG.

The high temperature part 11 and the low temperature part 1 of the cylindrical spacecraft 2.
A flat plate-shaped heat control panel 1 is provided along the axis so as to cover the solar rays 0 by crossing them. The surface of the thermal control panel 1 on the side of the spacecraft 2, that is, the surface (rear surface) opposite to the surface on the side of the sunlight 3 has a large absorptance a of the solar rays and a small emissivity ε. As a material having both of these properties, for example, there are those shown in Table 1.

[0013]

[Table 1] As can be seen from Table 1, when a galvanized Al plate is used, for example, a = 0.94 and ε = 0.1 are obtained.

In the above, the sun rays 3 hit the surface of the heat control panel 1 on the sun ray side and are shielded, and the high temperature portion 11 is not heated.

On the other hand, the albedo 5 from the earth 6 hits the back surface of the thermal control panel 1. Since the solar light absorption rate a is large, it is absorbed there. In addition, the earth radiation 4 from the earth 6 also hits the back surface. Since the thermal emissivity ε is small, it is reflected there and hits the high temperature portion 11 and the low temperature portion 10 to be heated.

Therefore, in the case where the heat control panel 1 is not provided, the high temperature part 11 and the low temperature part 1 which are generated by the sun rays 3 are generated.
0 means that the sun rays 3 are blocked and the earth radiation 4 hits,
Always weakly heated. Therefore, it is heated uniformly.

In this way, it is possible to prevent a partial positional and temporal temperature difference of the spacecraft 2 from being generated by the sun rays 3.

Further, the thermal control panel 1 does not need mechanical strength / rigidity, as long as it can optically block / reflect / absorb. Therefore, the degree of freedom in attaching to the spacecraft 2 is great, and the volume and weight at launch can be reduced. (2) A second embodiment of the present invention will be described with reference to FIG.

The thermal control panel 1a of this embodiment is almost the same as that of the first embodiment, but has a curved surface type whose back surface is loose and is concave along the machine axis.

By doing so, the reflected earth radiation 4 is often concentrated on the high temperature portion 11 and the low temperature portion 10 of the spacecraft 2. Therefore, the efficiency is better. (3) A third embodiment of the present invention will be described with reference to FIGS.

The heat control panel 1b is made of cloth, and one side thereof is attached to the spacecraft 2 along the axis thereof. A frame 8 having one end hinged with a pin parallel to the machine axis direction is provided at a position separated by a predetermined distance along the circumferential direction. The other side of the thermal control panel 1b is attached to the other end of the frame 8.

In the above, as shown in FIG. 3, the frame 8 is launched in a state of falling to the spacecraft 2 side. Then, the frame 8 expands on the orbit, and the thermal control panel 1b becomes a flat surface as shown in FIG.

The operation is similar to that of the first embodiment. Although not shown, the low temperature side is also provided. (4) A fourth embodiment of the present invention will be described with reference to FIGS.

In this embodiment, a panel 1a which is substantially the same as the heat control panel of the second embodiment is used, its mounting end is hinge-coupled 9, and the extension mechanism 18 of the panel 1a is mounted. It folds as shown in FIG. 5 until it gets on the track, and unfolds as shown in FIG. 6 on the track. The operation is almost the same as above. (5) A fifth embodiment of the present invention will be described with reference to FIGS.

In this embodiment, a panel 1a which is substantially the same as the heat control panel of the second embodiment is used, and both ends of the spacecraft 2 along the circumferential direction are attached by wire-like extending mechanisms 19 and 20, respectively. It is a thing. It contracts as shown in FIG. 7 until it gets into the orbit, and unfolds as shown in FIG. 8 on the orbit. The operation is almost the same as above.

[0026]

As described above, according to the present invention, since the sun rays are shielded, the albedo from the earth is almost absorbed, and the earth radiation is mostly reflected to the spacecraft side,
The temperature difference between parts of the spacecraft can be greatly reduced. Therefore, it is possible to prevent the occurrence of defects due to the temperature difference of the spacecraft.

[Brief description of drawings]

FIG. 1 is an overall configuration diagram of a first embodiment of the present invention.

FIG. 2 is an overall configuration diagram of a second embodiment of the present invention.

FIG. 3 is a configuration diagram of a third embodiment of the present invention.

FIG. 4 is an explanatory view of the operation of the embodiment.

FIG. 5 is a configuration diagram of a fourth embodiment of the present invention.

FIG. 6 is an explanatory view of the operation of the embodiment.

FIG. 7 is a configuration diagram of a fifth embodiment of the present invention.

FIG. 8 is an explanatory view of the operation of the embodiment.

FIG. 9 is an overall configuration diagram of a conventional spacecraft.

[Explanation of symbols]

1, 1a, 1b Thermal control panel 2 Spacecraft 3 Sunbeam 4 Earth radiation 5 Albedo 6 Earth 7 Orbit 10 Part where cosmic heat radiation is hard to reach (low temperature part) 11 Part where sunlight is directly incident (high temperature part)

Claims (1)

[Claims] 1. A solar cell is installed at a position where it crosses the sun's rays to cover the high temperature part and the low temperature part of the spacecraft, shields the sun's rays, and has a large solar ray absorption rate on the surface of the spacecraft side and radiation on the same surface. Thermal control panel for space, characterized by low rate.