JPH0460363B2 - - Google Patents

Description

[Detailed Description of the Invention] [Industrial Application Field] The present invention relates to a deployable antenna reflector,
In particular, it relates to the configuration of a deployable antenna mounted on an artificial satellite or space station, for example.

[Conventional technology]

Figures 4 to 7 are, for example, JP-A-59-
A front view of a conventional deployable antenna reflector disclosed in Publication No. 28704 when it is stored, a side view when it is stored,
They are a front view after development and a side view after development. In each of the above figures, 11a is a hinge that has a built-in elastic return force such as a spiral spring that serves as the driving force for rotation, and is equipped with a latch device that locks the rotation when a predetermined rotation angle is reached. Hinge A, 11b that latches at 11b is hinge B that also latches at a deployment angle of 135 degrees, and here, each hinge A
11a and the hinge B11b may be configured to incorporate a bearing capable of achieving low friction conditions, and this bearing may be, for example, a spherical bearing. 12
a, 12b are each hinge A11a, hinge B11b
These are tubular members whose ends are connected to each other to form a hoop shape as a whole, for example, a tubular long frame and a short frame made of carbon fiber composite material, and the long frame 12a is the satellite main body shown in FIG. The short frame 12b forms a hoop shape. Reference numeral 13 denotes a long frame 12a and a large number of support wires attached to each hinge A11a and hinge B11b, which are held with adjustable tension. Reference numeral 14 denotes two mesh-shaped flexible thin films whose edges are supported by a large number of support wires 13, and at least one of which is conductive. Reference numeral 15 refers to two opposing flexible thin films 14 that are moved inwardly toward each other. A tensile bonding wire, the length of which is adjustable. 16 is a spacer for holding each long frame 12a and short frame 12b during storage;
Reference numeral 7 denotes a fixture for fixing the long frame 12a to the satellite main body 18.

Next, the operation of the conventional deployable antenna reflector described above will be explained. At the time of launch, as shown in Figures 4 and 5, the ends are attached to each hinge A11.
a, each long frame 1 connected by hinge B11b
2a and short frame 12b is folded at each of the hinges A11a and B11b. Here, the long frame 12a is attached to the satellite body 18 via the fixture 17,
and a long frame 12 located opposite to this
As shown in FIG. 4, the length a is slightly longer than twice the length of the other short frame 12b in order to prevent the left and right folded hinges A11a from colliding with each other. Furthermore, two mesh-shaped flexible thin films 14 are provided inside the hoop via a large number of support wires 13 and are connected to each other by a large number of bonding wires 15.
It is folded between 2b. In this state,
The spacers 16 provided on the adjacent long and short frames 12a and 12b are in contact with each other and are fixed to the satellite main body 18 while being tightly bound by an appropriate fastening device (not shown). After reaching the orbit, when the binding device is released, each hinge A11a,
Each length and short frame 12 is adjusted by the torque of an elastic spring such as a spiral spring built into the hinge B11b.
a, 12b begin to unfold, and along with this, the two flexible thin films 14 folded between the long and short frames 12a, 12b also begin to unfold. Here, hinge A that was folded inward when stored.
11a is locked by the built-in latch device when the unfolding angle reaches 180 degrees, and the hinge B11b, which was folded outward, is locked by the built-in latch device when the unfolding angle reaches 135 degrees. . Thus, when all the hinges A11a and B11b are locked, an octagonal hoop is finally formed as shown in FIG. In this state, the length of each support wire 15 is adjusted in advance so that a predetermined tension is applied to the two mesh-shaped flexible thin films 14 connected by a large number of bonding wires 15. In addition, the length of each bonding wire 15 is set so that one flexible thin film 14 forms a parabolic surface (strictly speaking, the bonding point between each bonding wire 15 and the flexible thin film 14 is on a parabolic surface). has been adjusted in advance according to its position.

[Problem that the invention seeks to solve]

Since the conventional deployable antenna reflector described above is configured as described above, when the reflector is fully deployed, all the hinges A1
1A, hinge B11b is locked and the reflector maintains its shape. In this state, for example, the artificial satellite injects thrusters (not shown) for attitude control and the like. This excites elastic vibrations in the antenna reflector, and if the frequency of these elastic vibrations becomes too low, interference will occur with the attitude control system, so it is necessary to keep the lowest natural frequency of the antenna reflector high. There is. Normally, in the lowest vibration mode, the frame deflection near the satellite attachment point is large, and in order to increase the frequency of this mode, it is necessary to relatively increase the rigidity of the satellite attachment point without increasing the weight. . However, in conventional deployable antenna reflectors, the shape of the frame (especially the cross-sectional shape) is uniform for all frames, so if the cross-section of the frame is uniformly enlarged in an attempt to increase the lowest natural frequency,
This problem resulted in an increase in weight and an increase in volume when stored, which was a fatal flaw for an artificial satellite payload that places importance on light weight and small volume.

The present invention has been made to solve these problems, and an object of the present invention is to obtain a deployable antenna reflector that has a high lowest natural frequency and is small in weight and storage volume.

[Means for solving problems]

In the deployable antenna reflector according to the present invention, the rigidity of the frame is changed for each frame, and the rigidity is highest near the satellite attachment point, and the rigidity decreases as the distance from the satellite attachment point increases.

[Effect]

In the deployable antenna reflector of this invention, the rigidity of the frame near the satellite attachment point, which receives the highest load in the lowest natural vibration mode, is made relatively high, and the rigidity of the frame far from the satellite attachment point, which receives almost no load, is relatively increased. This increases the lowest natural frequency without increasing overall weight.

〔Example〕

FIG. 1 is a front view of a deployable antenna reflector according to an embodiment of the present invention after being deployed, and FIGS. 2 and 3 are cross-sectional views of the frame of the deployable antenna reflector of FIG. 1. In each of the above figures, 11a is a hinge that incorporates an elastic spring such as a spiral spring that provides the driving force for rotation, and is equipped with a latch device that locks the rotation when a predetermined rotation angle is reached. hinge A,
Reference numeral 11b is a hinge B that also tilts at a deployment angle of 135 degrees, and here, each hinge A11a and hinge B11b can be configured to have a built-in bearing that can realize low friction conditions, and this bearing can be replaced with, for example, a spherical bearing. It can be done. 22a, 22b and 23
a, 23b, 23c are tubular members whose ends are connected to each other by each hinge A11a and hinge B11b and form a hoop shape as a whole, for example, a tubular long frame and a short frame made of carbon fiber composite material, Each long frame 22a, 22b
are provided at an attachment portion to the satellite main body 18 and at a position opposite thereto on the hoop, and the other short frames 23a, 23b, and 23c form a hoop shape. 13 each long frame 22a,
22b and a number of support wires mounted on each hinge A11a, hinge B11b, with adjustable tension. 14 are two mesh-shaped flexible thin films whose edges are supported by a large number of support wires 13, and at least one of which is conductive;
Reference numeral 15 denotes a bonding wire that pulls the two facing flexible thin films 14 inwardly, and the length thereof can be set to be adjustable.
Reference numeral 17 denotes a fixture for fixing the long frame 22a to the satellite main body 18.

Next, the operation of the deployable antenna reflector which is an embodiment of the invention described above will be explained. After arriving in orbit from the stored state at launch,
Each hinge A11a and hinge B11b are expanded by built-in elastic springs, and the angle of the hinge A11a is adjusted.
When the angle of the hinge B11b reaches 180 degrees and the angle of the hinge B11b reaches 135 degrees, the rotation is locked by the latch device and the shape is maintained, which is the same as the conventional deployable antenna reflector. Here, as shown in Figures 1 to 3, the hollow cross-sectional shape of the frame is
Change it depending on the distance from 8. That is, the cross-sectional dimensions of the five frames near the satellite main body 18 are made larger than those of the conventional example, and the cross-sectional dimensions of the five frames far from the satellite main body 18 are made smaller than those of the conventional example, The cross-sectional shapes of the middle four frames are the same as those of the conventional example. For example, the thickness of the hollow section is the same for all frames. The hollow section is square,
If the width of the cross-sectional shape is a, the thickness is t, the Young's modulus is E, and the specific mass is p, the bending rigidity Sb and the mass per unit length _p1 are given by the following equations.

Sb=E/12{a ⁴ −(a−2t) ⁴ }≒2/3Eta ³ p ₁ =P{a ² −(a−2t) ² ≒4pta Due to injection from the attitude control thruster in the satellite body 18, etc. The antenna reflector vibrates in a cantilever beam mode, the amplitude of which increases as the distance from the satellite main body 18 increases. At this time, the cross-sectional shape of the frame is increased near the satellite main body 18 which receives a relatively large bending load, and the satellite main body 1 does not receive much bending load.
By reducing the cross-sectional shape of the frame in the portion far from 8, the rigidity against the cantilever vibration mode can be increased. Since the bending stiffness Sb changes as the cube of the cross-sectional shape width a, this cross-sectional shape width a
Even a slight change in this can have a large effect. In addition, since the mass per unit length changes linearly with the cross-sectional width a, if the cross-sectional width a is reduced in the portion far from the satellite body 18 by the amount increased near the satellite body 18, the overall The weight increase can be almost eliminated. Conversely, if it has the same rigidity, it is possible to realize a lighter antenna reflector.

In the above embodiment, the cross-sectional dimensions of the frame were changed in three steps, but in addition to this, the cross-sectional dimensions may be changed in multiple steps, or they may be changed continuously within the same frame. The same effects as in the embodiment are achieved.

〔Effect of the invention〕

As explained above, the present invention has a deployable antenna reflector in which the rigidity of a large number of frame-shaped frames supporting a flexible thin film is made highest near the satellite attachment point, and increases as the rigidity increases away from the satellite attachment point. Since the structure has a structure in which the rigidity is low and varies from frame to frame, an antenna reflector with extremely high rigidity can be obtained, and when the rigidity is the same, a lightweight antenna reflector can be obtained, which is an excellent effect. be.

[Brief explanation of the drawing]

FIG. 1 is a front view of a deployable antenna reflector according to an embodiment of the present invention after deployment, FIGS. 2 and 3 are sectional views of the frame of the deployable antenna reflector of FIG. 1, and FIGS. Figure 7 is a front view of a conventional deployable antenna reflector when it is retracted.
They are a side view when stored, a front view after deployment, and a side view after deployment. In the figure, 11a...hinge A, 11b...
Hinge B, 12a, 22a, 22b... long frame, 12b, 23a, 23b, 23c... short frame, 13... support wire, 14... flexible thin film, 15... bonding wire, 16... spacer, 1
7... Attachment, 18... Satellite main body. In addition,
In each figure, the same reference numerals indicate the same or corresponding parts.

Claims (1)

[Claims] 1. A large number of hinges each having a rigid spring that provides a driving force for rotation and a latch device that locks the rotation when a predetermined rotation angle is reached, and each hinge allows the ends to be connected to each other. A number of frames are connected together to form a hoop shape as a whole, and two mesh-shaped sheets are surrounded by each frame and supported at their edges by the frames and hinges via support wires, at least one of which is conductive. A flexible thin film is disposed between these two flexible thin films, and the two opposing points on the flexible thin film are pulled inward to form a desired shape as an antenna reflector. A deployable antenna reflector comprising a number of bonding wires applied to two flexible membranes, and a fixture disposed on one of the frames and serving as a coupling to the satellite body, wherein each frame has a cross-sectional shape. A deployable antenna reflector characterized in that the rigidity of the deployable antenna reflector is changed by decreasing the distance from the fixture. 2. The deployable antenna reflector according to claim 1, wherein each of the frames is made of a tubular member, and this member is made of a carbon fiber composite material. 3. The deployable antenna reflector according to claim 1 or 2, wherein the length of the coupling wire is adjustable. 4. The deployable antenna reflector according to claims 1 to 3, wherein the support wire is held with adjustable tension. 5. The deployable antenna reflector according to claims 1 to 4, wherein the hinge has a built-in bearing that can realize low friction conditions. 6. The deployable antenna reflector according to claim 5, wherein the bearing built into the hinge is a spherical bearing.