JPS5820599A - Artificial satellite - 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 spin-stabilized artificial satellite having a solar array paddle.

Conventionally, in artificial satellites that perform vertical attitude stabilization, multiple solar cells are combined electrically and attached to the outer surface of the satellite, and the power generated is supplied to onboard equipment. Ru.

However, according to this method, the solar receiving area is determined by the satellite shape, and as shown in Figure 1, the solar receiving area is determined by the spin axis + Il
In the case of a cylindrical artificial satellite (2) that spins around the cylindrical part, even if a solar cell array (3) is attached to the outer surface of the cylindrical part, the sunlight receiving area is less than 1/w of the outer surface of the cylindrical part. . Further, since the attachment area also depends on the outer surface area, the power obtained from the solar cell array (3) is limited by one dimension of the satellite shape. There are two methods that can be considered to improve the limit of the sunlight receiving area of the solar cell array.

The first method is to place a panel with a solar cell array attached on the outside of the outer circumference of the artificial satellite, as shown in Figure 2, and to move the panel along the outer circumference of the artificial valves in orbit. This is a method of increasing the sunlight receiving area. That is, in FIG. 2, +41 is a hollow cylinder with a solar cell array (abJ) attached to the outer surface of the above panel, and (2) is the satellite body with a solar cell array (3a) attached to the outer surface. (2) and ;4) both rotate around the spin axis (1).

Although this method can increase the area where the solar cells are attached, it is not possible to improve the ratio of the sunlight receiving area to the area where the solar cells are attached.

- 2nd method - Three-axis attitude stabilization method that controls the attitude of the satellite by rotating the wheels mounted on the satellite body rather than by spinning the satellite body. in.

This is a method of obtaining electricity from a solar array paddle attached to a solar array that is deployed from the satellite itself while in orbit.

With this method, the area where solar cells are attached does not necessarily depend on the outer circumferential area of the satellite.

In addition, the ratio of the sunlight receiving area to the solar IE pond area is also controlled by the solar array paddle.

This can be improved by orienting the solar cell attachment surface toward the sun. However, with this method, compared to the attitude stabilization method using Subi/, not only does the attitude control device for the 9 satellites become more complicated, but also the attitude stability of the 8 satellites is inferior, and the thermal control of the 1 satellite becomes more complicated.

This invention eliminates the drawbacks of the above nine methods.
While taking advantage of posture stability through spin.

By despinning the deployable thick battery panel against the spinning satellite body, it increases the ratio of the sunlight receiving area to the solar cell attachment area, and eases restrictions on the solar cell attachment area for one satellite shape. The drawings will be described in detail below.

FIG. 3 shows an embodiment of the present invention, and shows nine artificial satellite bodies (
2) Spin rate signal (6) and solar direction reference signal + from attitude sensors such as the sun sensor and earth sensor mounted on the
71 '&- is input to the despan motor drive control circuit (5) to control the rotational speed and phase of the despan motor (8), and is attached to the rotor side of the despan motor f81.
Control is performed so that the solar cell attachment surfaces of the iI battery paddles (3a) and (3b) face the sun. The electric power generated by the solar battery paddles (4a) and (4b) is transferred to the slip ring (9a) and (
9b) Derived to the Fuyuyama 9 satellite body (2).

FIG. 4 shows an example of the external appearance of an embodiment of the present invention, in which semi-cylindrical solar panels (4a) and (4b) are attached to the satellite main body (2) that slides around the spin axis il+.
The despan section attached to the atma rotates in the opposite direction to the satellite main body, and is controlled so that the solar cell attachment surfaces of the solar cell panels (4a) and (4b) face toward the sun.

Hinge parts (, 1la), (llb), (llc) and (lid) and mounting shafts (12aJ and (12bJ)
is used to fold the solar panels (4a) and (x) around the satellite body for better storage when one satellite is stored in the fairing of the launch vehicle during the satellite launch stage. 1 When the satellite is launched into orbit, it is expanded and the hinge part (11a),
(llb), (llc) and (lid) are latched.

In addition, in FIG. 4, the case where the artificial satellite main body has a cylindrical shape has been explained, but there are no restrictions on the shape of one artificial satellite main body. Furthermore, although the number and shape of the unfolded solar battery paddles have been described using two semi-cylindrical panels, there are no restrictions on the number or shape. Especially the satellite itself is 4
In the case of a prismatic solar cell panel, by using a flat solar cell panel and folding the solar cell panel multiple times, it is possible to create a large solar cell within the constraints of the fairing dimensions of the rocket. The advantage is that the battery panel can be stored.

Figure 5 shows an example of implementation on an artificial satellite equipped with a directional communication antenna circle mounted on the Despan Platform aJ. Then, the umbilical line α5) is routed and supplied to the satellite body (2). Thick 11#3
Despan section 1 fl (with battery panel +4) attached
j and the satellite main body (2) The despan part is made by fully fixing the clamp! Rotation of tl [) becomes possible, and at the same time, the umbilical line (151) is also disconnected by the clamp and the slow-acting mechanical switch (16). Solar array attached to (
3) supplies power to the equipment mounted on the satellite main body (2), and the power generated by the main battery panel 141 is transferred to the slip ring (91) and the tea towel despan platform (
Despan section +21) Q3KM mounted equipment Q? + supplies power.

In order to direct the despan part zastcN-controlled communication antenna α4 toward the earth station, the despan motor drive circuit (5) is used to rotate the despan part 2 (13) in the opposite direction to the satellite main body (2). Elide span motor 2
(8b) Control IJ. The despan motor drive circuit (5) is connected to the earth sensor mounted on the satellite body 12+, and the sun'
A spin rate signal (6) from an attitude sensor such as a sensor,
It receives the earth direction reference signal and the sun direction reference signal (7), and fully controls the rotational speed and phase of the despan section 1 (11m) and the despan section 203.

The relative angular velocity between the despan section I IIG and the despan section 2 (131) is the orbital angular velocity, that is, 360°A for a satellite in a geostationary orbit.
A) is simpler than a high-speed slip ring and has improved reliability. Also.

Despan section 2 (Equipment mounted on 13 (17+ and slip ring (9b) on the opening of the satellite body (2) no longer needs to include a power line/ This also makes it possible to achieve optical coupling by means of a directional antenna (141), thereby increasing reliability. Fig. 6 is an example of an external appearance when the present invention is implemented on a despan platform type artificial satellite having a directional antenna (141). In Figure 6, (2) is the main body of the satellite that spins around the spin axis +11, and 001 is the solar panel (4a) and (X) with the solar cells attached so that the surface faces the sun. (lla), (llb), (11cJ
and (11dJ is the solar panel (4m) and (4b
) Deployment latch! ! It is an I structure hinge, (1
2a) and (12b) are solar panels (4a) and (
4b) is the mounting shaft. The old 1 is the despan section 2 that carries communication equipment, and the despan section 2 (u4) is a directional antenna attached to the despan section 2 (13).The despan section 2t131 spins so that the directional antenna circle is directed toward the earth station. The satellite body (2) rotates around the axis in the opposite direction at approximately the same speed as the satellite body (2).The solar array (3) rotates in the opposite direction to the satellite body (2) after the solar battery panels (4a) and (4b) are deployed. When this invention is implemented in the despan platform type artificial satellite described above, even if the strength requirements of the equipment mounted on the despan platform part are changed,
Despan platform section and solar panel'f!
: Just by changing the two satellites itself (this can be applied).

As described above, in the artificial satellite according to the present invention.

By despinning the deployable solar panel against the spinning satellite body, it is possible to expand the sunlight receiving area and increase the light receiving efficiency of the solar cell attached surface compared to conventional spin-stabilized satellites. ^, it can be said that it is an artificial satellite suitable for carrying equipment that requires a large amount of electric power.

[Brief explanation of the drawing]

Figures 1 and 2 are external views of a conventional spin-stabilized artificial satellite, Figure 3 is a configuration diagram of an embodiment of this invention, Figure 4 is an external view of an embodiment of this invention, and Figure 5 is an external view of a conventional spin-stabilized artificial satellite. FIG. 6 is a block diagram of another embodiment of the invention. FIG. 6 is an external view of another embodiment of the invention. In the figure, (11 is the spin axis, (2) is the spinning satellite body, (13) is the solar cell array, +4) is the solar cell panel, (
5) is the despan motor drive circuit, (61 is the spin rate signal, (7) is the solar direction base IS sound signal 18) is the despan motor, (9) is the slip ring, aO is the solar panel mounting despan part, uu #'i Hinge with latch mechanism. α21 is the solar panel mounting shaft, Q31u despan platform part, (141 is the directional antenna, [1
51 is an umbilical line, α61 is a clamp interlocking mechanical switch, α71 is a device mounted on the despan platform, and 081 is an earth direction reference signal. Note that the same or equivalent parts are indicated by the same reference numerals. Agent Shin Kuzuno - Figure 111 "1 Figure 112 Figure 1 Figure 113

Claims (1)

[Claims] In a satellite that has a deployable solar panel with a solar array attached to it and is configured to stabilize its attitude through spin, the solar battery panel is attached to the satellite body in a despin (direction of spin direction of the satellite). An artificial satellite characterized in that the surface of the solar cell panel to which the solar cells are attached is directed toward the sun by rotating the solar cell panel in the opposite direction.