WO2016087721A1

WO2016087721A1 - Method for manufacturing an antenna reflector shell, in particular of a space vehicle

Info

Publication number: WO2016087721A1
Application number: PCT/FR2015/000214
Authority: WO
Inventors: Audrey-Marine LOUIS
Original assignee: Airbus Defence And Space Sas
Priority date: 2014-12-03
Filing date: 2015-11-25
Publication date: 2016-06-09
Also published as: FR3029690A1; FR3029690B1

Abstract

The invention relates to a method for manufacturing an antenna reflector shell (1), wherein said shell (1) includes a sandwich structure (4), provided with a central element (5) and two skins (6, 7) arranged on either side of the central element (5) and including one skin (6) referred to as front skin, said front skin (6) reflecting electromagnetic radiation and being made of a composite material. Said method includes a perforation step that involves, during the production of the shell (1), perforating the shell (1) such as to make a plurality of holes therein that pass entirely through the shell (1).

Description

A method of manufacturing an antenna reflector shell, particularly a

^Spacecraft

The present invention relates to a method of manufacturing an antenna reflector shell, in particular a spacecraft.

Although not exclusively, the present invention applies more particularly to a hull forming part of an antenna reflector of a spacecraft such as a telecommunication satellite, in particular a large antenna reflector. .

It is known that antenna reflectors generally comprise a shell comprising a rigid surface, which is reflective (for certain electromagnetic waves), and reinforcing elements (or structure) provided at the rear of the shell, which participate in maintaining the hull and connecting with the satellite.

This antenna reflector geometry can be used in a wide range of frequency bands, between 1GHz and 60GHz (L, S, C, X, Ku, Ka, Q, V bands).

Antenna reflectors must meet strict specifications concerning both the reflectivity of electromagnetic waves on the reflecting surface and the mechanical strength of the entire reflector (shell and structure) to the mechanical and acoustic stresses induced by space launchers. .

The shell of an antenna reflector is generally composed of a sandwich-type structure formed of a central honeycomb element on which are affixed a front skin and a rear skin. Each of these skins consists of one or more plies (or layers) of composite material provided with fibers, in particular carbon fibers. Each fold can be a unidirectional fold or a woven fold. The choice of materials constituting the front skin (ie the reflective surface) of the hull shall to ensure proper reflection of electromagnetic waves.

The large size of the reflectors and thus of the hull induces significant mechanical stresses on the reflector during the launch of the satellite. These mechanical stresses are, in particular, due to acoustic stresses (drum skin effect). Acoustic forces may correspond to the most critical cases of mechanical dimensioning of the reflector.

For this type of reflector, the stresses induced by the acoustic stresses become much greater than the stresses induced by the other mechanical stresses.

The object of the present invention is to remedy this drawback and to enable the manufacture of a large reflector shell which is insensitive to acoustic stresses, and for which the stresses induced by the other existing stresses are greater than the acoustic stresses. More precisely, it seeks to reduce the impact of acoustic demands, without increasing the impact of other solicitations.

To do this, the present invention relates to a method of manufacturing an antenna reflector shell, in particular an antenna of a spacecraft, said shell comprising a sandwich-type structure, provided with a central element ( or soul) and two skins arranged on either side of said central element and comprising a so-called skin before skin, said front skin being reflective for electromagnetic radiation and being made of a composite material.

According to the invention, said manufacturing method comprises a perforation step consisting, during manufacture of the shell, to perforate the shell so as to practice a plurality of holes completely through the shell. Thus, thanks to the invention, by producing a pierced shell, that is to say pierced with numerous holes, of small size as specified below, the impact of the acoustic stresses on the shell of the shell is reduced. reflector. Thus, the drumskin effect is greatly diminished because the acoustic waves (ie the air pressure waves) can pass through the reflector shell. In addition, the impact of acoustic stress is reduced, without increasing the impact of other solicitations.

In addition, a perforated shell allows the dissipation of a portion of the forces applied to the shell by viscous damping of the air or by resonance in the thickness of a shell comprising a central element consisting for example of a nest of bees, which also reduces the impact of the efforts on the reflector.

In a preferred embodiment, the central element (preferably honeycomb) of the shell is provided with through openings, and the perforation of the shell consists of producing, successively, the perforation of the one and the other of said skins. It is also conceivable to perform at the same time the perforation of the two skins.

In a first (preferred) embodiment, the perforation (during the perforation step) is performed using at least one laser.

In this first embodiment:

the perforation step consists in making a relative displacement between the laser and the shell to be perforated; and or

the perforation consists in providing and maintaining a constant spacing between an external surface of one of said skins to be perforated and the laser, this gap being substantially equal to the focal length of the laser.

Furthermore, in a second embodiment, the perforation is performed using at least one mechanical tool.

In addition, advantageously:

the perforation is carried out after a polymerization of the composite material at least from the front skin; and or

- The perforation is to practice holes of round section in said skins; and or

perforation consists in making holes distributed in said skins in a given regular distribution. Preferably, the regular distribution of the holes is adapted to the arrangement of fibers in the composite material constituting the corresponding skin; and or

the perforation consists in making elongated holes having a longitudinal general direction, the longitudinal general direction of each of said holes being substantially normal to an external surface of the corresponding skin; and or

- The perforation is implemented automatically according to a preset program.

The present invention also relates to an antenna reflector shell, in particular a spacecraft antenna, obtained by implementing the aforementioned method, that is to say a perforated shell, said perforated shell comprising a sandwich-type structure, provided with a central element and two skins arranged on either side of said central element and comprising a so-called skin before skin, said front skin being reflective for electromagnetic radiation and being made of a composite material.

The present invention further relates to:

an antenna reflector, in particular an antenna of a spacecraft, comprising a shell as mentioned above; and

a spacecraft, in particular a satellite, comprises at least one such antenna reflector.

The appended figures will make it clear how the invention can be realized. In these figures, identical references designate similar elements. Figure 1 is a partial schematic sectional view of an antenna reflector shell.

Figure 2 is a schematic perspective view of a sandwich structure of a shell.

Figure 3 is a schematic plan view of a portion of a central element of a shell.

Figure 4 is a schematic partial view, in perspective, of a portion of a perforated skin of a shell.

Figure 5 is a partial view, in section, schematically showing a perforation of the shell.

FIG. 6 schematically shows a device for implementing a perforation step during a method of manufacturing a shell.

The present invention relates to a method of manufacturing a shell 1 of an antenna reflector, in particular an antenna of a spacecraft, as shown schematically in FIG.

More particularly, although not exclusively, this shell may be part of an antenna reflector of a spacecraft such as a telecommunications satellite, in particular a large antenna reflector.

Such an antenna reflector comprises, generally:

a shell 1 representing a rigid structure (or panel) provided with a reflective surface on one. face 2 called front face; and

- Reinforcing elements (not shown) which are arranged at a rear face 3 of the shell 1, opposite the front face 2. These reinforcing elements participate in the maintenance of the shell 1 and the connection with the satellite. These reinforcing elements are known and not described further in the present description.

In a preferred embodiment, the shell 1 comprises a structure 4 of the sandwich type. As shown in FIG. 2, this structure 4 is provided with a central element (or core) 5, specified hereinafter, and two skins 6 and 7 arranged on either side of said central element 5, namely:

a so-called front skin 6 which is arranged at the front face 2 of the shell 1 and which is able to receive and send back the electromagnetic radiation received by the antenna reflector; and

- A so-called rear skin 7 which is arranged at the rear face 3 of the shell 1 and which is intended in particular to receive the reinforcing elements.

The front skin 6 is therefore configured to be reflective for electromagnetic radiation having frequencies adapted to the frequency band or bands intended to be reflected by the reflector, and it is made of a composite material.

In a preferred embodiment, the rear skin 7 is also made of a composite material.

Preferably, the sandwich-type structure 4 comprises a central element (or core) 5 of the NIDA type (structural and transparent to electromagnetic waves) reinforced by the two skins 6 and 7, preferably prepregs of carbon fibers / epoxy matrix. The NIDA (ie honeycomb) structure is provided with a plurality of hexagonal shaped cells 8, as shown in FIG. 3, which makes it possible to reinforce the resistance of the shell 1 while guaranteeing maximum lightness.

For reasons of simplification of the drawing, the structure 4 is represented in rectangular general form in FIG. 2. Of course, in the case of an antenna reflector shell, this structure 4 has a generally circular shape, for example with a diameter of the order of two to three meters, and it is curved in cross section (as shown in Figure 1).

The general steps of the manufacturing process of the shell 1, except a perforation step specified below, which are usual steps known to those skilled in the art and adapted in particular to the constituent material killing the structure 4 of the shell 1, are not described further in the present description. By way of illustration, a particular customary manufacturing process can make it possible to manufacture the shell with composite material based on carbon fiber and resin. It can use, for this, unidirectional folds or fabrics that are draped on a mold and polymerized.

According to the invention, the method of manufacturing the shell 1 comprises at least one perforation step of perforating the shell 1 so as to practice a plurality of through holes 9, as shown in part in Figure 4.

The central element 5 of the shell 1 is provided with through openings, in particular cells 8 (FIG. 3) in the case of a honeycomb structure, and the perforation of the shell 1 (implemented during of this perforation step) is to successively perform the perforation of one and the other of said skins 6 and 7, preferably by first perforating one of the skins (for example the skin 6) and then returning the structure 4 to perforate the other skin (eg skin 7) with the same device or tool. Figure 4 schematically illustrates a portion of a skin 6, 7 and perforated.

It is also conceivable to achieve the perforations of the two skins 6 and 7 at the same time (namely a pass per side, or a pass for both sides).

Therefore, by manufacturing a perforated shell 1, that is to say, pierced with many holes, of smaller dimension as indicated below, the impact of the acoustic stresses on the shell 1 is reduced. Drum-skin effect is greatly diminished because acoustic waves (ie, air pressure waves) can pass through shell 1, as illustrated by an arrow F in Figure 5. Acoustic waves pass through the holes 9 of the skin before 6, then the openings 8 of the central element 5, and finally the holes 9 of the rear skin 7. To do this, it is not necessary that the holes 9 of the skin 6 and 7 (and in particular D1 and D2) are aligned, as is the case in the particular example of Figure 5.

In addition, the impact of acoustic stresses is thus reduced without increasing the impact of other stresses, particularly without adding reinforcing carbon composite plies on the reflector (which would weigh down the reflector).

In addition, a perforated shell 1 makes it possible to dissipate part of the forces applied to the shell 1 by viscous damping of the air or by resonance in the thickness of the central element 5 (preferably consisting of a nest of bees), which also reduces the impact of the efforts on the reflector.

The perforation is performed after polymerization of the composite material (consisting of fibers and resin) of the skins 6 and 7.

In a first embodiment, the perforation of the shell 1 is performed using at least one conventional mechanical tool (not shown), for example using a drill.

Furthermore, in a second preferred embodiment, the perforation of the shell 1 is carried out using at least one laser 10, as shown in FIG. 6. This laser 10 is part of a perforation device 11 .

Thanks to the use of the laser 10, a particularly rapid method is obtained for perforating the shell 1.

In this second preferred embodiment, the perforation step consists in implementing a relative displacement between the laser 10 and the shell 1 to be perforated in a plane P (substantially parallel to the plane of the shell).

This displacement is implemented by appropriate means, in particular mechanical means 12 or optical means. The mechanical means 12 may comprise a set of guides 13 and a set of motor means 14, for example one or more electric motors. 3 in the example of FIG. 6, the laser 10 is moved above the shell 1 in one direction and one direction, illustrated by an arrow E.

Upstream of the current position of the laser 10 (along the direction E), the holes 9 are already practiced. The perforation is performed by the emission by the laser 10 of a laser beam 15.

The perforation of the shell 1 consists in successively perforating the one and the other of said skins 6 and 7 with the aid of the laser 10, by first perforating one of the skins (for example the skin 6 as shown in FIG. 6) located at the top, facing the laser 10, then rotating the structure 4 by 180 ° to bring the other skin (for example the skin 7) in the correct position, facing the laser 10, for to be able to perforate this other skin.

Laser perforation consists in locally subliming the material. Thus, the fibers (in particular of carbon) of the skins 6 and 7 made of composite material do not comprise (or little) of broken or misaligned fibers, which minimizes the impacts of passive intermodulation product (or PIM, that is, that is, nonlinearities in the reflection of electromagnetic waves introduced by discontinuities). In addition, the surface state of the reflective zone (face 2 of the shell 1) is not modified, and the geometry of the reflective surface is not impacted.

The laser perforation is adapted to an embodiment of holes 9 of dimension (for example diameter for holes of round section) less than 0.5 mm. It is known that the use of antenna reflectors in high frequency bands corresponds to the use of waves whose wavelength is relatively low. The waves are sensitive to geometries (and in particular to holes 9) whose size has the same order of magnitude as the wavelength. So for reflective applications antennas in the frequency bands considered (bands L to V for frequencies below 60 GHz), the laser perforation is able, unlike other (mechanical) means, to make holes 9 adapted in diameter. less than 0.5 mm.

The laser perforation 10 allows perforations to be made according to a specifically selectable hole size, hole shape and hole distribution.

In a particular embodiment, during perforation, the device 11 maintains a constant spacing D between the external surface (front face 2) of the skin 6 in question, and the laser 10. This spacing is preferably substantially equal to the focal length of the laser 10.

The use of a laser 10 for perforating has other advantages:

the holes 9 obtained are very regular. In addition, the shape of the holes 9 is controlled with great precision (within a few micrometers);

the holes 9 may have sections of different shapes. The preferred form of the holes is a round section, as shown in Figure 4. The latter form allows the fastest perforation.

The distribution of the holes 9 in the plane of the skin 6, 7 can also be selected. For example, the holes 9 may be aligned and distributed according to a given distribution, for example square (as shown in Figure 4), rectangular, triangular or other. The holes 9 can also be arranged individually without repeating a particular pattern.

The preferred distribution of the holes is an alignment of the holes 9 with the direction of the fibers (carbon for example) in the composite material of the skin 6, 7 concerned, to perform the perforation along some fibers only. Thus, a minimum of fibers are cut, which guarantees good mechanical performance of the shell 1. For example, a distribution of holes in squares (Figure 4) is adapted to the use of a hull whose skins are biaxial tissues. The size and the density of the holes 9 are chosen so as to allow good air circulation (as a function of the thickness of the air boundary layer on the shell).

Moreover, thanks to the laser perforation, all the holes 9 made have a uniform surface state for each hole 9 and between the holes 9.

In a preferred embodiment, the perforation can be performed industrially during the manufacture of reflector shells. In this case, the laser 10 is adapted to a tool (device 11) which makes it possible to move the laser 10 over the entire surface of the shell 1, with a high accuracy and a good capacity for repetition of the positioning of the holes.

The tooling is configured to maintain a constant distance D between the laser 10 and the skin surface 6, 7. Depending on the settings and the choice of the laser 10, this distance D may vary, but as indicated above, the preferred distance between the laser 10 and the surface to be perforated is equal to the focal length chosen for the laser 0.

The perforation device 11 also comprises mechanical or optical means (not shown) for orienting the laser beam 15 with respect to the surface to be perforated. The laser beam 15 is preferably oriented in a direction normally normal to the surface of the shell, for better efficiency.

More specifically, in a preferred embodiment, the perforation consists in making elongated holes having a longitudinal general direction, namely, as shown in the example of FIG. 5, a longitudinal general direction D1 for a hole 9 of the skin 6 and a longitudinal general direction D2 for a hole 9 of the skin 7. The longitudinal general direction D1, D2 of each of the holes 9 is substantially normal to the external surface (towards the outside) of the skin 6, 7 considered (face 2, 3). The choice of the type of laser and the adjustments to be applied depend on the material constituting the skin 6, 7 of the shell 1, the thickness of the skin 6, 7, and the geometry of perforation that one wishes to achieve. In particular, it is possible to select the power of the laser 10, the focal length, the scanning speeds, the cutting strategy of each hole, and the order of perforation of the holes.

In a preferred embodiment, the perforation device 11 comprises a control unit 17 for controlling the laser 10 according to the parameters and the strategy determined in advance. To do this, the control unit 17 is linked via a link 18 to the motor means 14 and via a connection 19 to the laser 10. These parameters and strategy are programmed, in the usual manner, in the control unit 17 and automated.

Note that the shell 1 can be held in place on a specific tool or on the mold that allowed to drape. The laser 10 can be installed in a landscaped area near the manufacturing zone of the shell, which makes it possible subsequently to implement the construction of the shell and the perforation of the shell, in an industrially effective manner.

In addition, the manufacturing method, as described above, does not introduce dissymmetry in the composite stack constituting the shell 1. Dissymmetries have indeed very negative impacts on the thermomechanical behavior of the reflector subject to variations. high temperatures.

The realization of an antenna reflector insensitive to acoustics is also of industrial interest. It is not necessary to systematically perform an acoustic stress test to achieve the reflector qualification, which generates a gain in terms of cost and planning.

Claims

1. A method of manufacturing an antenna reflector shell, in particular an antenna of a spacecraft, said shell (1) comprising a structure (4) of sandwich type, provided with a central element (5). ) and two skins (6, 7) arranged on either side of said central element (5) and comprising a so-called skin front skin (6), said front skin (6) being reflective for electromagnetic radiation and being produced in a composite material,

characterized in that it comprises a perforation step consisting, during the manufacture of the shell (1), to perforate the shell (1) so as to practice a plurality of holes completely through the shell (1), and that, the central element (5) of the shell (1) being provided with through openings, the perforation of the shell (1) consists in perforating said two skins (6, 7).

2. Method according to claim 1,

characterized in that the perforation of the shell (1) consists in producing, successively, the perforation of one and the other of said skins (6, 7).

3. Method according to one of claims 1 and 2,

characterized in that the perforation is performed after polymerization of the composite material at least the front skin (6).

4. Method according to one of claims 1 to 3,

characterized in that the perforation is performed using at least one laser (10).

5. Method according to claim 4,

characterized in that the step of perforating is to achieve a relative displacement between the laser (10) and the shell (1) to be perforated.

6. Method according to one of claims 4 and 5,

characterized in that the perforation comprises providing and maintaining a constant spacing (D) between an outer surface (2) of one (6) of said skins to perforate and the laser (10), this spacing (D) being substantially equal to the focal length of the laser (10).

7. Method according to one of claims 1 to 3,

characterized in that the perforation is performed using at least one mechanical tool.

8. Method according to any one of the preceding claims, characterized in that the perforation consists in making holes (9) of round section in said skins (6, 7).

9. Method according to any one of the preceding claims, characterized in that the perforation consists in making holes (9) distributed in said skins (6, 7) in a given regular distribution.

10. Process according to claim 9,

characterized in that the regular distribution of the holes (9) is adapted to the arrangement of fibers in the composite material constituting the corresponding skin.

11. Method according to any one of the preceding claims, characterized in that the perforation consists in making holes (9) elongated having a longitudinal general direction (D1, D2), the longitudinal general direction (D1, D2) of each of said holes (9) being substantially normal to an outer surface of the skin (6, 7) thereof.

12. Method according to any one of the preceding claims, characterized in that the perforation is implemented automatically according to a preset program.

13. Antenna reflector shell, in particular of a spacecraft antenna,

characterized in that it is obtained by carrying out the method specified in one of claims 1 to 12.

Antenna reflector, in particular of a spacecraft antenna, characterized in that it comprises a shell (1) according to claim 13.

15. Spacecraft, in particular a satellite,

characterized in that it comprises at least one antenna reflector according to claim 14.