Structural component, aircraft or spacecraft, and method

Abstract

A structural component for an aircraft or spacecraft, comprising: a planar member; a reinforcing member which projects from the planar member and is rigidly connected thereto; the reinforcing member comprising at least a foam layer and a cover layer, a plurality of pins extending at least through the foam layer and the cover layer, and at least the pins and the cover layer comprising a curable matrix.

Claims

1. A structural component for an aircraft or spacecraft, comprising: a planar member; and a reinforcing member which projects from the planar member and is rigidly connected thereto; wherein the reinforcing member comprises at least a foam layer and a cover layer, wherein a plurality of pins individually extend continuously through the planar member, the foam layer, and the cover layer, and wherein at least the pins and the cover layer comprise a curable matrix.

2. The structural component of claim 1, wherein the foam layer or the pins or the foam layer and the pins are connected directly to the planar member.

3. The structural component of claim 2, wherein the foam layer or the pins or the foam layer and the pins are glued directly to the planar member.

4. The structural component of claim 1, wherein the reinforcing member comprises a further cover layer, which is arranged between the planar member and the foam layer and connected thereto, the further cover layer likewise comprising a curable matrix.

5. The structural component of claim 4, wherein the pins extend through the further cover layer.

6. The structural component of claim 4, wherein at least one of the cover layer, the further cover layer, the planar member or the pins comprise a fibrous material, which is infiltrated by the respective matrix.

7. The structural component of claim 6, wherein the fibrous material comprises at least one of an interlaid scrim, woven fabrics or rovings.

8. The structural component of claim 6, wherein two of the reinforcing members are provided, and cross at a crossing point, and at the crossing points, fibres of the fibrous material which extend in the longitudinal direction of the respective reinforcing members are exclusively those which extend over the entire length of a respective reinforcing member.

9. The structural component of claim 1, wherein the planar member is formed as a skin portion.

10. The structural component of claim 1, wherein a plurality of the reinforcing members are provided and together form a grid structure.

11. The structural component of claim 10, wherein a plurality of the reinforcing members together form a diamond structure.

12. An aircraft or spacecraft comprising a structural component of claim 1.

13. A method for manufacturing a structural component, comprising: providing a planar member; providing a reinforcing member which projects from the planar member and is rigidly connected thereto, wherein the reinforcing member comprises at least a cover layer and a foam layer; introducing a plurality of pins, wherein the plurality of pins individually extend continuously through the planar member, the foam layer, and the cover layer; arranging the foam layer, the cover layer and the pins on the planar member so as to form the reinforcing member projecting from the planar member; infiltrating at least the cover layer and the pins with a curable matrix and curing the matrix; and connecting the reinforcing member to the planar member.

14. The method of claim 13, wherein the pins are introduced by at least one of inserting prefabricated rigid pins, stitching in fibres or incorporating fibres at least into the foam layer and the cover layer.

15. The method of claim 13, wherein a further cover layer is provided between the foam layer and the planar member and connected thereto.

16. The method of claim 13, wherein the pins are initially inserted at least into the foam layer and the cover layer and if applicable into the further cover layer, subsequently the foam layer or if applicable the further cover layer is connected to the planar member or the foam layer and the cover layer and if applicable the further cover layer are arranged on the planar member, and subsequently the pins are introduced into the foam layer, the cover layer, the planar member and if applicable the further cover layer.

17. The method of claim 13, wherein the cover layer and if applicable the further cover layer are formed from a fibrous material, in particular a fibrous interlaid scrim or fibrous woven fabrics, fibres of the fibrous material in particular extending exclusively in a respective longitudinal direction of two reinforcing members which are to be manufactured and parallel to this direction, triangular, diamond or rectangular regions being cut out of the fibrous material, longitudinal edges of the regions extending parallel to the respective longitudinal directions of the reinforcing members.

18. The method of claim 13, wherein the cover layer, the pins and if applicable the further cover layer are infiltrated with the curable matrix and cured at a first time, and the planar member is infiltrated with a curable matrix and cured at a second time, the first and second times being the same or different times.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The invention is described in greater detail in the following by way of embodiments, with reference to the appended drawings, in which:

(2) FIG. 1 is a perspective view of a detail of a structural component according to an embodiment of the present invention;

(3) FIG. 2 is an enlarged view A from FIG. 1;

(4) FIG. 3 is a section B-B from FIG. 2 in one state of a method according to an embodiment of the present invention;

(5) FIG. 4 is the section B-B of FIG. 3 in accordance with a variant B-B of the method; and

(6) FIG. 5 is a plan view of a fibrous material for manufacturing reinforcing members of the structural component of FIG. 1.

DETAILED DESCRIPTION OF THE INVENTION

(7) In the figures, like reference numerals denote like or functionally equivalent components unless otherwise stated.

(8) FIG. 1 is a perspective view of a detail of a structural component 1 of an aircraft generally denoted as 2.

(9) The structural component 1 comprises a planar member 3 and reinforcing members 4, 4. The reinforcing members 4, 4 are firmly attached to the planar member 3. The reinforcing members 4, 4 form a diamond structure having an angle see FIG. 2of between for example 15 and 45, preferably between 30 and 40. The angle is the acute angle of a respective diamond 5.

(10) FIG. 2 is an enlarged view A from FIG. 1.

(11) The reinforcing members 4, 4 project from the planar member 3. In other words, an underside 6, concealed in FIG. 2, of each of the reinforcing members 4, 4 lies on the planar member 3, and an upper side 7 thereof is spaced apart from the underside 6 or upper side of the planar member 3 by a height H. The height H may for example be between 25 and 35 mm. Further, the reinforcing members 4, 4 may have a width W of for example 35 to 70 mm. The height H and the width W relate in the above to a cross-section Q of a respective reinforcing member 4, 4. In the present embodiment, the cross-section Q is rectangular in shape. However, other cross-sections are equally conceivable. The length L of a respective reinforcing member 4, 4 can be anything from a few centimeters to several meters.

(12) FIG. 3 is a section B-B from FIG. 2 in one state of a method according to an embodiment of the present invention.

(13) It can be seen from FIG. 3 that the reinforcing member 4 is composed of an upper cover layer 11, a lower cover layer 12, and a foam layer 13 arranged between the cover layers 11, 12. The cover layers 11, 12 each comprise a fibrous material 14, described in greater detail below with reference to FIG. 5, which is embedded in an epoxy resin matrix. The foam layer 13 may in particular be a rigid foam.

(14) In the method steps preceding the state shown in FIG. 3, the cover layers 11, 12 are stitched to the foam layer 13 positioned between them by way of a stitching thread 15, as is shown highly schematically on the left-hand side of FIG. 3 by way of a dashed line. The stitching thread 15 is for example formed as roving, and therefore comprises a plurality of individual threads. The reinforcing member 4 formed in this manner is infiltrated with an epoxy resin matrix, the fibrous material 14 of the cover layers 11, 12 being infiltrated as well as the stitching thread 15. Subsequently, the reinforcing member formed in this manner is cured, forming stable pins 16. The pins 16 are of an elongate shape and may in particular be formed with a circular cross-section. The fibres in the respective pins 16 extend along the longitudinal direction L.sub.1 thereof. The longitudinal direction L.sub.1 of the respective pins 16 is preferably oblique to a normal S to the planar member 3, for example at an angle of between 10 and 80. The stitching preferably takes place in the longitudinal direction of a respective reinforcing member 4, 4, and this results in two parallel rows of pins 16, as can be seen particularly clearly from FIG. 2.

(15) Subsequently, the formed reinforcing member 3 is cured. To provide the reinforcing member 4 with a corresponding shape, it can be oriented on a correspondingly contoured laminated device before being cured. In this way, a curved shape of the reinforcing member 4 can for example be achieved.

(16) After the reinforcing member 4 is cured, it is applied to the planar member 3. The planar member 3, which is in particular a fuselage outer skin, preferably likewise comprises a fibrous material which is embedded in an epoxy resin material. The planar member 3 is thus formed for example as a prepreg. After the reinforcing member 4 is applied to the planar member 3, the planar member 3 is cured, whereupon the epoxy resin matrix in the planar member 3 connects it firmly to the reinforcing member 4.

(17) However, a number of variants are possible in this context. For example, the cover layers 11, 12 may comprise a prepreg. Accordingly, the pins 16 may also be formed with an epoxy resin matrix which has not yet been cured. The reinforcing member 4 which has not yet been cured may thus be applied to the planar member 3 while still flexible. In a subsequent step, the structural component 1 formed in this manner is cured as a whole. As a further alternative, the reinforcing member 4, which is provided with the matrix but has not yet been cured, could be placed on the planar member 3 which has already been cured, whereupon the structural component 1 formed in this manner is cured as a whole.

(18) FIG. 4 shows the section B-B of FIG. 3 in accordance with a variant B-B of the method.

(19) In this variant, the reinforcing member 4 is constructed on the planar member 3. For this purpose, the foam layer 13 is applied directly to the planar member 3. However, it would equally be possible also to apply the lower cover layer 12 to the planar member 3 beforehand, and only subsequently to apply the foam layer 13 to the lower cover layer 12. Subsequently, the upper cover layer 11 is applied to the foam layer 13. Subsequently, the upper cover layer 11, the foam layer 13 and the planar member 3 are stitched together by means of the stitching thread 15. After a corresponding infiltration with an epoxy resin matrix, the pins 16 are formed. The configurations and variants which were described previously in connection with FIG. 3 can be applied correspondingly to the embodiment of FIG. 4.

(20) Instead of the stitching process, it is further conceivable to use a different type of process, in which the pins 16 are prefabricated so as already to be stable, and are subsequently inserted into the composite consisting of the upper cover layer 11, the foam layer 13 and the planar member 3. The prefabricated pins 16 are thus already cured. After the pins 16 are inserted, the epoxy resin matrix of the upper cover layer 11 of the planar member 3 and if applicable the lower cover layer 12 are subsequently cured.

(21) The pins 16 may moreover be formed by incorporating the fibrous material 15 into at least the foam layer 13 and the upper cover layer 11 by means of a needle, as disclosed in DE 10 2005 024 408 A1, of which the content to this effect is also the subject-matter of the present application.

(22) In the embodiments of FIGS. 4 and 5, the pins 16 may each extend through the upper cover layer 11, the lower cover layer 12 and the planar member 3, as is shown for two pins 16 in FIG. 4.

(23) FIG. 5 is a plan view of the fibrous material 14 for forming the grid-like reinforcing members 4, 4.

(24) The fibrous material 14 is composed of fibres 21, 22. The fibres 21 and 22 extend at the angle to one anothersee FIGS. 2 and 5. The fibres 21 and 22 may be interwoven. Alternatively, the fibres 21 and 22 may be part of different interlaid scrims which are positioned one above the other. The fibres 21 and 22 initially form a planar formation, for example an approximately rectangular formation. FIG. 5 only shows parts of this formation.

(25) Subsequently, the diamond-shaped regions 5 are cut out from the formationsee FIGS. 2 and 5. For this purpose, the fibres 21 and 22, shown in dashed lines in FIG. 5, are cut to length, in such a way that merely the fibres 21 and 22 shown as a solid line in FIG. 5 remain. In FIGS. 2 and 5, the longitudinal direction of the reinforcing member 4 is denoted by the reference numeral L.sub.2 and the longitudinal direction of the reinforcing member 4 is denoted by the reference numeral L.sub.3. The cutting lines 26 and 27 parallel to the longitudinal directions L.sub.2 and L.sub.3, for cutting out a respective diamond-shaped region 5, are shown as a dot-dash line in FIG. 5 for merely one of the regions 5.

(26) In a plan view, the fibrous material 14 produced in this manner corresponds to the progression of the reinforcing member 4, 4see FIG. 2. The fibrous material 14 is distinguished in that there are no fibres doubled up at the crossing points 23see FIGS. 2 and 5between the fibres 24, associated with the reinforcing member 4 and extending in the longitudinal direction L.sub.2 thereof, and the fibres 25, associated with the reinforcing member 4 and extending in the longitudinal direction L.sub.3 thereof.

(27) The fibrous material 14 may subsequently be used to form the upper cover layer 11 and if applicable the lower cover layer 12. The fibre formation 14 is infiltrated with an epoxy resin matrix and subsequently used as described in connection with FIGS. 3 and 4.

(28) The presently used fibres are preferably carbon fibres, although inclusions of other fibres such as aramid fibres are also conceivable. Whenever reference is made to an epoxy resin matrix in the present document, this may also comprise a proportion of thermoplastic.

(29) Although the invention has been disclosed by way of preferred embodiments, it is not limited thereto, but can be modified in various ways. The embodiments and developments disclosed for the structural component according to the invention apply correspondingly to the aircraft or spacecraft according to the invention and to the method according to the invention, and vice versa. It should further be noted that in the present document, the term a does not exclude a plurality.