Method for manufacturing an aircraft structure component

Abstract

A method for manufacturing an aircraft structure component of fiber composite material, wherein a semifinished first part made of a dry fiber preform is provided, a semifinished second part made of either a dry fiber preform or a hardened fiber matrix composite is provided, a semifinished aircraft structure component is prepared by placing the semifinished first part against the semifinished second part, such that a first connection surface rests against a second connection surface, matrix material is infused through the semifinished aircraft structure component, and the semifinished aircraft structure component is cured. Before infusing matrix material through the semifinished aircraft structure component, an intermediate layer for assisting adhesion is applied between the first connection surface and the second connection surface.

Claims

1. A method for manufacturing an aircraft structure component of fiber composite material having a first part and a second part connected to the first part, the method comprising the steps of: a) providing a semifinished first part made of a dry fiber preform and having a first connection surface, b) providing a semifinished second part made of at least one of a dry fiber preform and a hardened fiber matrix composite, wherein the semifinished second part has a second connection surface for being connected to the first connection surface of the semifinished first part, c) preparing a semifinished aircraft structure component by placing the semifinished first part against the semifinished second part, such that the first connection surface rests against the second connection surface, d) infusing matrix material through the semifinished aircraft structure component, and, e) curing the semifinished aircraft structure component, wherein before infusing matrix material through the semifinished aircraft structure component, an intermediate layer for assisting adhesion is applied between the first connection surface of the semifinished first part and the second connection surface of the semifinished second part.

2. The method according to claim 1, wherein the intermediate layer is formed as an adhesive film.

3. The method according to claim 2, wherein the adhesive film is applied between the first connection surface and the second connection surface by placing the adhesive film solidly, integrally, or both onto the first connection surface or the second connection surface.

4. The method according to claim 2, wherein the adhesive film is applied between the first connection surface and the second connection surface by spraying an aerosol onto at least one of the first connection surface and the second connection surface.

5. The method according to claim 2, wherein the adhesive film is applied between the first connection surface and the second connection surface by distributing at least one of a liquid and a viscous resin onto at least one of the first connection surface and the second connection surface.

6. The method according to claim 2, wherein the adhesive film is perforated, comprises a structured surface, or both.

7. The method according to claim 1, wherein the intermediate layer is formed as at least one of a fleece of dry and preimpregnated fibers.

8. The method according to claim 1, wherein the intermediate layer is formed as at least one of a fabric of dry and preimpregnated fibers.

9. The method according to claim 1, wherein one of the semifinished first part and the semifinished second part is formed as a semifinished skin element, and wherein the other one of the semifinished first part and the semifinished second part is formed as a semifinished support element.

10. The method according to claim 9, wherein the semifinished first part is formed as the semifinished skin element and the semifinished second part is formed as the semifinished support element.

11. The method according to claim 10, wherein the semifinished skin element comprises an inner surface including the first connection surface, and wherein the semifinished support element comprises a flange portion including the second connection surface.

12. The method according to claim 1, wherein after the intermediate layer has been applied and before the matrix material is infused, the intermediate layer is pre-cured.

13. The method according to claim 1, wherein the intermediate layer, after being applied, is fixed to at least one of the first connection surface and the second connection surface by fixing means.

14. The method according to claim 1, wherein the semifinished first part is placed on a tool and the semifinished second part is placed on the semifinished first part.

15. The method according to claim 14, wherein for curing, the tool together with the semifinished aircraft structure component placed thereon, is exposed to at least one of a certain temperature in an oven, and a certain pressure and a certain temperature in an autoclave.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0027] In the following, a preferred embodiment of the present invention is explained in more detail by means of a drawing. The drawing shows in:

[0028] FIG. 1 a cross sectional view of an aircraft structure component manufactured by a method according to an embodiment of the present invention.

DETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION

[0029] As shown in FIG. 1, an aircraft structure component 1 of fiber composite material having a first part 3 and a second part 5 connected to the first part 3, can be manufactured by a method according to the invention. In the present embodiment, the aircraft structure component 1 relates to a fuselage shell segment 7, wherein the first part 3 corresponds to a skin element 9 and the second part 5 corresponds to a support element 11, in particular a stringer element, for supporting the skin element 9. The method comprises the following steps:

[0030] First, a semifinished first part 3 and a semifinished second part 5 are provided. Both the semifinished first part 3 and the semifinished second part 5 are made of a dry fiber preform. The semifinished first part 3 has a first connection surface 13 and the semifinished second part 5 has a second connection surface 15 for being connected to the first connection surface 13. The semifinished first part 3 is formed as a semifinished skin element 9 which comprises an inner surface 17 including the first connection surface 13. The semifinished second part 5 is formed as a semifinished support element 11 which comprises a flange portion 19 including the second connection surface 15.

[0031] Subsequently, an intermediate layer 21 for assisting adhesion is applied on the first connection surface 13 of the semifinished first part 3 or on the second connection surface 15 of the semifinished second part 5. The intermediate layer 21 in the present embodiment is formed as an adhesive film 23, in particular a FM 300 film of Cytec Industries, which is applied on the first connection surface 13 or on the second connection surface 15 in a solid, integral form, in particular by unwinding it from a roll and placing it on either the first connection surface 13 or the second connection surface 15. In the present embodiment the adhesive film 23 is perforated and comprises a structured surface for allowing and promoting gas exchange of the underlying first or second connection surface 13, 15.

[0032] As a next step, a semifinished aircraft structure component 1 is prepared by placing the semifinished first part 3 against the semifinished second part 5, such that the first connection surface 13 rests against the second connection surface 15. In such a way, the adhesive film 23 is sandwiched between the semifinished skin element 9 and the flange portion 19 of the semifinished support element 11, in particular between the first connection surface 13 and the second connection surface 15. The semifinished aircraft structure component 1 is positioned on a corresponding tool 25 for forming and supporting the desired shape of the aircraft structure component 1 to be manufactured, wherein the semifinished first part 3 is placed on the tool 25 and the semifinished second part 5 is placed on the semifinished first part 3. In such a way, the outer surface 27 of the semifinished skin element 9 opposite to the inner surface 17 rests on the tool 25.

[0033] Then, the surface 29 of the semifinished aircraft structure component 1, in particular the surface 29 of the semifinished aircraft structure component 1 which is not in contact with the tool 25, is covered by a vacuum bag 31.

[0034] Subsequently, the semifinished aircraft structure component 1 together with the tool 25 on which it is placed is positioned in an autoclave (not shown) or an oven (not shown). In the autoclave or oven the semifinished aircraft structure component 1 is preheated up to an infusion temperature, so that the adhesive film 23 is pre-cured between the first and second connection surfaces 13, 15 in order not to be washed away by infused matrix material 33.

[0035] At the infusion temperature matrix material 33 is infused through the semifinished aircraft structure component 1, i.e. through and between the semifinished first part 3 and the semifinished second part 5, by applying a vacuum 35 to the environment under the vacuum bag 31.

[0036] After the matrix material 33 has been entirely infused, the semifinished aircraft structure component 1 is cured by applying a certain pressure and a certain temperature in the autoclave over a certain time, or by applying only a certain temperature in the oven over a certain time.

[0037] Finally, after curing is completed, the finished aircraft structure component 1 can be taken out of the autoclave or oven. The intermediate layer 21, i.e. the adhesive film 23, considerably strengthens the connection between the support element 11 and the skin element 9, so that the flange portion 19 of the support element 11 can be minimized for a minimum weight of the aircraft structure component 1.

[0038] As is apparent from the foregoing specification, the invention is susceptible of being embodied with various alterations and modifications which may differ particularly from those that have been described in the preceding specification and description. It should be understood that I wish to embody within the scope of the patent warranted hereon all such modifications as reasonably and properly come within the scope of my contribution to the art.