JPS6130497A - Hinge mechanism of expanding structure of space missile - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a deployment device for a spacecraft, and particularly to a hinge mechanism for a deployment structure such as a solar array paddle or an antenna mounted on an artificial satellite.

The solar array paddles and antennas onboard the satellite are housed within the fairing of the launch vehicle, so for example,
As shown in the figure, the solar panel 3 is folded and stored on the side of the satellite body 2 and is launched, and after the satellite body 2 is separated from the rocket, the panel 3 is opened and the dashed line 3' in the figure It is common for the structure to be expanded as shown in . In addition, 1 is a rocket fairing, 4
is a hinge portion that connects the panel 3 so that it can be opened and closed.

FIGS. 2(a) and 2(b) show a plan view and a side view of the object to be developed, that is, the solar cell panel 3, in the unfolded state, using a solar cell paddle as an example of the deployable structure.

Two rigid hinge parts 4 are fixed to each side of each solar panel 3, and rigid shafts 5 are inserted into the bearing holes of these hinge parts 4 independently from each other to form a hinge mechanism. has been done. By the way, in recent years, artificial satellites have become larger, and as a result, deployable structures have also become larger.Therefore, there is a tendency for deployable structures to be required to be lighter and have more compact storage space. Regarding the panels to which solar cells are attached, instead of the conventional rigid honeycomb board, there is a blanket type as shown in Figures 3(a) and 3(b), in which the outer periphery of a thin film 6 such as Kapton membrane is reinforced with a composite material 7. flexible panels 8.

However, if you attach two or three conventional independent and rigid hinge mechanisms to such a blanket-type panel 8 and try to deploy it, excessive stress will be concentrated near the hinge mechanism attachment parts and cracks will occur. do. If a highly rigid reinforcing plate is attached as a measure to prevent cracks in the blanket, the weight will increase and the purpose of using a thin film blanket type will be lost. Furthermore, if the number of hinge connections between the blankets is small, the shape of the blanket will be distorted as shown in FIG. 4 when the blankets are deployed, and the directions of the hinge shafts will be uneven, making it impossible to deploy the blankets smoothly. Furthermore, connecting a large number of conventional hinge mechanisms to the blanket in order to eliminate this drawback has the drawback of increasing weight.

The present invention was made in order to eliminate the above-mentioned drawbacks of the conventional technology, and by constructing the hinge connection part from a composite material such as fiber, the hinge connection part is made compact and lightweight, and further developed. It is an object of the present invention to provide a hinge mechanism for a spacecraft deployable structure in which stress concentration of an object to be deployed in the vicinity of a hinge attachment portion that occurs therein is dispersed.

The present invention also includes: By making the hinge shafts of each hinge connection part one instead of independent, it eliminates irregularities in the axis direction of the shafts and allows the entire hinge part to change according to the shaft displacement curve. The purpose of the present invention is to provide a hinge mechanism for a spacecraft deployment structure that enables deployment.

Hereinafter, embodiments of the present invention will be described with reference to the drawings.

5(a) is a plan view of a solar cell panel to which the present invention is applied, and FIG. 5(b) is a side view of the panel shown in FIG. 5(a). Further, FIGS. 6(IL) and 6(b) are enlarged side views of a hinge mechanism portion according to an embodiment of the present invention. Note that Figures 5(a) and (b) show one panel taken out, and Figures 6(a) and (
b) shows a state in which the panels are connected by a hinge mechanism and unfolded by 180° IC in the direction of the arrow. As shown in FIGS. 5(a) and 5(b), the outer periphery of a thin film 6 such as a Kapton membrane is covered with a composite material 7 such as carbon fiber to form a flexible panel 8 of a blanket tibe. A plurality of edges of the composite material 70 are stretched out laterally in the form of tongues at opposite sides of the panel, and the stretched portions of the brackets are folded back into hollow bead shapes, thereby forming a plurality of edges of the composite material 70. Spaced bearing portions 9 are formed. In one embodiment, the thickness of the composite material (carbon fiber, etc.) of the bearing portion 9 is 0.3
fi, and the folded end portion is fixed to the composite material 7 on the outer periphery of the thin film body 6 by adhesive or the like. With this structure, a plurality of panels 8 are joined together so that the bearing parts protruding from the sides are alternately assembled, and the bearing holes are aligned. Shaft 10 (made of composite material) (Fig. 6 (λ
), see (b)) and connect it to the pressure. 6(a) (j) In the embodiment, the composite material constituting the bearing portion 9 is formed separately from the composite material 7 serving as the thin film reinforcement material, and the composite material 7 is bonded to the composite material 7. There is. Figure 6 (
The embodiment b) has a structure in which the composite material '2 and the t+q++ receiving part 9 are integrally molded as described above. The bearing part 9 is the fifth
As in the embodiment shown in Figure (a), a large number of hinge shafts are formed, and one hinge shaft 10 of the connecting portion is passed through in common. Adjacent panels 8 are shown in FIG. 6(a).

It unfolds around the shaft 10 of <4M portion as shown by the arrow in (b). With this structure, the hinge part can be made compact and lightweight, and the composite material 7.
Stress generated in the thin film body 6 is dispersed, and the orientation of the shaft does not change individually and unevenly, so smooth deployment is possible. In addition, since the structure is simpler than that of the conventional unfolding mechanism, the number of manufacturing steps can be greatly reduced, and there is no need to strictly define the accuracy of bearing mounting, resulting in a corresponding reduction in costs.

As explained in the above explanations 1 to 1, according to the present invention, the deployment mechanism is made of a composite material such as carbon fiber, so it is lightweight, compact, and easy to manufacture, and it is possible to smoothly deploy a thin-film solar panel. It has the advantage of being In addition, in the above-mentioned embodiment, explanation was given using a solar cell panel as an example.
It is clear that the configuration of the present invention can be applied to the deployment mechanism of antennas and other deployable items mounted on artificial satellites.Also, in the above embodiment, a case where one shaft is passed through in common is explained, but it is equivalent to this. The shaft structure may have the following effects, and it is clear that these modifications are also included in the present invention.

[Brief explanation of the drawing]

Figure 1 is a front view of the satellite body in the conventional solar array paddle storage state, Figures 2 (a) and (b) are a plan view and side view of the conventional deployment structure, respectively, and Figure 3 (a).
, (b) are respectively a plan view and a side view of a flexible solar melon pond panel, FIG. 4 is a perspective view of a conventional flexible deployable structure in a state where the axes of the hinge shafts are irregularly displaced, and FIG. Figures (a) and (b) are a plan view and a side view of a flexible solar panel according to an embodiment of the present invention, respectively;
FIGS. 6(a) and 6(b) are side views of the hinge portion of the flexible deployable structure according to the embodiment of the present invention. 1... Rocket fairing, 2... Satellite body, 3, 3'... Panel, 4
...Hinge part, 5...Rigid hinge shaft. 6...Thin film body, 7...Fabric material, 8...
Flexible panel, 9... Bearing portion, 10... Hinge shaft.

Claims (1)

[Claims] A hinge mechanism characterized in that a hinge connection part of a deployable structure of a spacecraft is formed of a composite material such as carbon fiber.