JPH1034825A - Bonding structure with low thermal strain of navigating body in space - Google Patents

Abstract

PROBLEM TO BE SOLVED: To obtain a bonding structure with low thermal strain of a navigating body in space with enough small error and deviation of a sensor orientation axis by arranging a metal layer with a plus linear expansion coefficient and a CFRP laminated collar with a minus linear expansion coefficient on the outside and the inside of a zero thermal strain truss. SOLUTION: The outermost layer of a bonded part is a metal layer 1 and a zero thermal strain truss 3 being a part to be bonded is bonded on the inside of the metal layer. In addition, a CFRP laminated collar 2 is bonded on the inside of the zero thermal strain truss 3. The zero thermal strain truss 3 is a main part constituting a truss structure and the truss part can be realized by a zero thermal strain CFRP laminated constitution. As the bonded part is an integral bonded structure, the metal layer 1 with a plus linear expansion coefficient and the CFRP laminated collar 2 with a minus linear expansion coefficient cancel each other their deformations to temp. change.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a low thermal strain bonding structure for a space vehicle such as an earth observation satellite.

[0002]

2. Description of the Related Art Many earth observation satellite structures and precision observation sensor structures are made of low heat distortion materials such as carbon fiber reinforced plastic (CFRP) in order to minimize the sensor pointing axis error in orbit. Have been. The joint between such low heat strain materials is made of an Al alloy or a Ti alloy.

FIG. 3 is a view showing a conventional joint, in which a metal joint 5 having a positive linear expansion coefficient and a CFRP material 6 having a negative linear expansion coefficient are joined in series to reduce heat distortion as a whole. I was aiming.

[0004]

A first problem of the prior art is that there is a limit to the reduction in thermal strain. The reason is that the temperature of each part of the spacecraft does not fluctuate uniformly, but the temperature distribution becomes non-uniform due to shade, sunshine, internal heat generation and the like.

[0005] In recent years, the performance of earth observation satellites and precision observation sensors has been further enhanced (high resolution), and the need to minimize sensor pointing axis errors has increased. On the other hand, artificial satellites have also been increasing in size and power, and there has been a tendency for sensor pointing axis fluctuations due to thermal strain to increase. However, in the related art shown in FIG. 3, the sensor pointing axis error and the sensor pointing axis fluctuation cannot be sufficiently suppressed.

Accordingly, an object of the present invention is to provide a low-heat-strain joint structure for a spacecraft in which thermal distortion can be greatly suppressed as compared with the conventional art, and as a result, sensor pointing axis error and sensor pointing axis fluctuation can be sufficiently reduced. To provide.

[0007]

According to the present invention, there is provided a space vehicle in which a metal layer having a positive linear expansion coefficient and a CFRP laminated collar having a negative linear expansion coefficient are arranged on the outside and inside of a zero thermal strain truss, respectively. Is obtained.

According to the low thermal strain bonding structure of the present invention, since the metal layer and the CFRP laminated collar cancel each other in response to a temperature change, a low thermal strain bonding structure can be obtained.

[0009]

Next, an embodiment of the present invention will be described. FIG. 1 shows an observation satellite structure employing the low thermal strain bonding structure of the present invention, and is an example of a truss structure bonding. FIG. 2 is a partially detailed view of the joint.

The structure shown in FIG. 1 includes a zero-heat-strain truss 3 and a truss joint 4. The truss joint 4 is a part for joining the zero-heat-strain truss 3 and includes a metal layer 1 and a CFRP laminated collar 2 as shown in FIG.

The outermost layer of the joining portion is a metal layer 1, and a zero-heat-strain truss 3, which is a portion to be joined, is adhered to the inside of the metal layer. Further, the CFRP laminated collar 2 is adhered to the inside of the zero heat strain truss 3 (see FIG. 2). Zero heat strain truss 3
The truss part is a main component of the truss structure shown in FIG.

Since the main joint has an integral bonding structure, the metal layer 1 having a positive linear expansion coefficient and the CFRP laminated collar 2 having a negative linear expansion coefficient are deformed by a temperature change. Cancel each other.

Assume that the cross-sectional area of the metal layer 1 is A1, the linear expansion coefficient is α1, the elastic coefficient is E1, the length is l, and the elongation with respect to the temperature difference Δt is Δl. Similarly, when the material properties of the CFRP laminated color 2 are A2, α2, E2, Δl = Δtα1 l (1) Δl = Pl / EA1 (2) From equations (1) and (2), P = Α1 E1 A1 Δt, P is
This is the thermal strain force generated in the metal layer 1 due to the temperature difference Δt. In order to balance this thermal strain force P with that of the CFRP laminated collar 2, the specifications are such that α1 E1 A1 = α2 E2 A2 (3) Set.

If the parameters of α, E, and A of the metal layer 1 and the CFRP laminated collar 2 are set by the formula (3), and the coefficient of linear expansion α of the joint is calculated as a specific example, in the case of an Al alloy, α1 = 22 × 10-6 / ℃ to α
= 1 × 10−7 / ℃ In the case of Ti alloy, α1 = 7 × 10−6 / ℃
About 1 × 10 −7 / ° C. is feasible.

With such a configuration, a structure with zero thermal strain in the length direction at any temperature can be realized, and a zero thermal strain structure of the entire satellite structure can be realized.

In the embodiment of the present invention, a truss structure, a metal material and a CFRP laminate have been described as examples, but it is clear that the present invention can be applied to all frame structures and other structures of a spacecraft. It is within the scope of the present invention.

It is also obvious that the present invention can be applied to the case where different materials having different linear expansion coefficients are joined (for example, ceramic and CFRP) to control the overall linear expansion coefficient. Within range.

[0018]

A first effect of the present invention is that zero thermal strain can be accurately realized. The reason is that the whole linear expansion coefficient calculation logic is obvious and the effect is obvious.

[Brief description of the drawings]

FIG. 1 is a diagram showing an observation satellite structure according to an embodiment of the present invention.

FIG. 2 is a partial detailed view of a joining portion in FIG. 1;

FIG. 3 is a diagram showing a conventional configuration.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Metal layer 2 CFRP color 3 Zero heat strain truss 4 Joint

Claims (3)

[Claims] 1. A space having an interlayer bonding structure of a metal layer and a CFRP laminate, wherein a metal layer having a positive linear expansion coefficient and a CFRP layer having a negative linear expansion coefficient cancel each other's expansion and contraction. Low thermal strain joint structure of the navigation body. 2. A joining structure for joining components constituting a space vehicle, wherein a metal layer is provided on the outside and a CFRP is provided on the inside.
A low thermal strain bonding structure for a spacecraft, wherein layers are arranged. 3. The low thermal strain bonding structure for a spacecraft according to claim 2, wherein the constituent elements of the spacecraft are zero heat strain trusses.