DE102013208215B4 - Contour-changing wing and method for producing such a wing - Google Patents

Abstract

Contour-changing wing (12), which extends in a span direction (S) and (i) a wing skin (14) and (ii) a contour changing device (16), by means of which a lift coefficient of the wing (12) by changing a contour of the wing skin ( 14) can be changed on a leading edge of the wing (10), characterized in that (iii) the wing skin (14) (a) a matrix (26) made of a matrix material, (b) at least one transverse stiffening body (28) which extends along extends in the span direction (S), and (c) has at least one longitudinal stiffening element (30) which extends along an inflow direction (A) of inflowing air during operation of the wing (12), (d) the matrix (26) being a has lower rigidity than the longitudinal stiffening element (30), so that when the contour of the wing skin (14) changes in a longitudinal bending plane (E) that is perpendicular to the transverse stiffening bodies (28), a neutral fiber (N) in the longitudinal stiffening element ment (30) runs.

Description

The invention relates to a contour-changing wing according to the preamble of claim 1. According to a second aspect, the invention relates to a method for producing such a wing. According to a further aspect, the invention relates to a fiber composite component.

In order to change the lift coefficient for wings, slats are known, for example Fowler flaps or nose flaps. If these flaps are actuated, the contour of the wing changes and thus the lift. A particularly high effect with considerably less noise development can be expected from contour-changing wings whose wing front edges can be changed in their contour. Another way to increase lift on aircraft is to provide a high lift flap on the trailing edge of the wing.

The problem with such contour-changing wings is that the leading edge of the wing must on the one hand absorb the considerable forces that act on it during flight, and on the other hand must be deformable in such a way that the contour can be changed sufficiently strongly. It has been found that with materials available hitherto, a sufficiently high deformability can only be achieved with material thicknesses below 1 millimeter. However, this leads to a low mechanical stability of the leading edge of the wing and the risk that, for example, the impact of a bird makes the wing at least partially unusable.

The US 2006/0 145 031 A1 describes an aircraft wing which has high rigidity in the span direction and high flexibility in the direction perpendicular thereto. In the span direction, stiffening elements made of carbon fiber reinforced plastic are present, which are arranged in the troughs of a corrugated sheet-like carbon fiber composite panel. The remaining volume of the aircraft wing is filled with an elastic material. In some embodiments, cavities are also present in the aircraft wing.

The WO 2009/118 509 A1 describes a fiber composite material that can be used, among other things, in the aerospace industry. The material has an increased electrical conductivity compared to conventional composite materials and consists of a resin in which carbon fibers are embedded in the longitudinal and transverse directions, and of a surface layer which comprises metal-coated fibers.

The DE 10 2009 003 696 A1 describes a membrane, in particular a gas storage membrane, which consists of at least two individual membrane sheets which are connected to one another by at least one seam construction, each individual membrane sheet containing at least one elastomer layer and at least one outer layer based on PTFE. Furthermore, a method for producing such a membrane is described.

The invention has for its object to reduce the noise development in wings.

The invention solves the problem by a contour-changing wing with the features of claim 1.

According to a second aspect, the invention solves the problem by a method having the features of claim 6. A cold-curing resin is understood to mean a resin which cures at 30 °, in particular 40 ° C., to 70 ° C., in particular 60 ° C. .

The invention also solves the problem by a method having the features of claim 8. This method is carried out in particular in an autoclave process with additional pressure (4-7 bar).

According to a further aspect, the invention achieves the object by means of a fiber composite component having the features of claim 9. If the fiber composite component is a wing that has a skin structure in the form of the fiber composite component on its leading edge of the wing, the transverse bending stiffness of the skin structure is one Bending in the span direction of the wing is at least ten times more rigid than bending the skin structure in the flow direction of the wing.

The fact that the fibers run at least essentially parallel to one another means that an arrangement parallel in the mathematical sense is admittedly cheap, but is neither achievable nor necessary. It is possible, for example, that at most 10% of the fibers have a deviation of more than 10 ° from the strict parallelism.

The advantage of such a contour-changing wing is that its wing front edge can be changed in its contour over a wide range and that at the same time the bending stiffness in the transverse direction is so great that the loads occurring can be safely absorbed by the incoming air. In other words, a distance from struts that absorb bending loads in the transverse direction and divert them into a wing base body can be chosen so large that their weight does not make the wing excessively heavy.

It is a further advantage that such a contour-changing wing can be produced using technically relatively simple means. It is also favorable that the leading edge of the wing can be easily assembled. The advantage of contour variability therefore does not have to be bought with a significantly higher weight.

The large achievable ratio of the stiffness in the span direction to the stiffness in the inflow direction of the wing is particularly advantageous. This stiffness ratio can be well over 10. Previously known wing skins do not even come close to this value. As a result, the actuators for deforming the leading edge of the wing can be made weaker and lighter. It is also advantageous that the wing skin shows no or only a slight anticlastic curvature due to the strong anisotropy.

In the context of the present description, a wing is understood to mean in particular an aircraft wing, that is to say a component of an aircraft, which is designed to generate lift when the aircraft is operating. The wing can also be a component of a wind turbine.

The contour changing device is understood in particular to be a device by means of which the contour of the leading edge of the wing can be changed in an adjustable, reproducible manner such that the lift coefficient changes.

The feature that the lift coefficient of the wing can be changed means in particular that the contour can be changed so that an effective curvature of the wing changes. In particular, the contour changing device is therefore designed in such a way that the wing can be brought into an elongated, non-curved or slightly curved arrangement and a more curved arrangement.

A change in the contour of the wing skin in the longitudinal bending plane is understood to mean that with such a change in contour, the curvature changes in the longitudinal bending plane, but not in a plane perpendicular to this longitudinal bending plane.

The matrix material is preferably an elastomer, for example a thermoplastic or a thermoset. If the sash is to be used at very different temperatures, the use of a thermoset elastomer has advantages. The elastomer is preferably rubber. For example, the AA6CFZ mixture from Kraiburg can be used.

According to a preferred embodiment, the longitudinal stiffening element forms a continuous, curved surface which extends along at least a portion of the leading edge of the wing. This has the advantage that the pressure that the air flowing towards the leading edge of the wing exerts on the leading edge of the wing can be absorbed by the longitudinal stiffening element at all points of impact.

It is advantageous if the longitudinal stiffening element has a thickness of at most 2 millimeters, in particular at most 1 millimeter. Such a longitudinal stiffening element can be deformed particularly easily and is nevertheless sufficiently stable to absorb the forces which the inflowing air exerts on the leading edge of the wing.

According to a preferred embodiment, the cross stiffening body is constructed from a fiber composite material and has a plurality of fibers running in parallel in a cross stiffening body matrix. In particular, the wing skin comprises a large number of transverse stiffening bodies, at least a large majority of which are made up of fiber bundles, each of which is accommodated in a transverse stiffening body matrix. The individual transverse stiffening bodies are disjoint from one another, that is to say separated from one another and from the longitudinal stiffening element by means of matrix material.

The material from which the cross-stiffening body matrix is constructed generally differs from the matrix material of the matrix. The material from which the cross-stiffening body matrix is constructed is selected such that the cross-stiffening body has a particularly high bending stiffness. For example, an epoxy resin is used for the cross stiffener matrix. The matrix material, on the other hand, is selected so that it is sufficiently firm on the one hand and sufficiently elastic on the other hand to allow a slight change in the contour. For example, rubber is used.

The material from which the cross-stiffening body matrix is constructed is therefore preferably chosen as a thermosetting resin. It is also advantageous if the material for the cross stiffening body matrix on the one hand and the matrix material on the other hand are selected such that they form a firm connection with one another which is so firm that bending of the wing skin does not lead to a detachment of the cross stiffening body from the surrounding matrix material.

The statements made above for the transverse stiffening body apply alternatively or additionally to the longitudinal stiffening element. It is possible, but not necessary, that the Cross stiffening body and the longitudinal stiffening element are constructed from the same fibers and the same resin.

In the context of a method according to the invention, the step of surrounding the fiber bundles with the matrix of elastomer is achieved, for example, by selecting a thermoplastic elastomer, for example thermoplastic polyurethanes, and by arranging this elastomer around prefixed fiber bundles. In a subsequent step, the matrix material is then heated so that it cross-links and encloses the fiber bundle. It is important, however, that the matrix material does not penetrate the fiber bundle, or at least only insignificantly. Simultaneously or in a separate step, the fibers of the fiber bundles are connected to one another by means of a resin, so that the cross-stiffening body matrix and / or the longitudinal stiffening element matrix is formed.

According to a preferred embodiment, the consolidation of the fibers comprises the steps of applying a pressure difference between a first end of the covered fiber bundle and a second end of the covered fiber bundle, which is opposite the first end, and of sucking and / or pushing the resin through the fiber bundle. In other words, the matrix is first formed, which surrounds the fiber bundle or possibly partially penetrates into the outermost layers of the fiber bundle. However, the inside of the fiber bundle remains free of matrix material. These gaps are then filled with the resin in the subsequent step. This resin is then cured so that it combines with the surrounding matrix.

An alternative method is carried out by simultaneously heating the fiber bundles (prepreg) impregnated with resin and the matrix material surrounding the fiber bundles, so that firstly the fibers of the fiber bundle combine with the resin, secondly the resin combines with the matrix material and thirdly the matrix material is cured to the matrix. In particular, curing is crosslinking, just as the resin is preferably cured by crosslinking.

In the first production process, it is advantageous if, in order to suck the resin through the dry fiber bundles, a vacuum of, for example, 0.1 to 0.8, in particular 0.6 to 0.8, for example 0.7 bar is applied. In addition, the resin can be injected from the opposite side at a pressure of 0.3 to 2.0, for example 1.1 bar.

The temperature at which this process is carried out depends on the materials used. If, for example, a thermoplastic polyurethane, for example polyurethane (ether) 4201 AU from Bayer, is used as the matrix, a temperature of 172 to 173 ° with a holding time of 10 to 15 minutes is advantageous. The most suitable temperature is determined in preliminary tests. A temperature that is too low leads to incomplete or slow wrapping of the fiber bundles of the cross-stiffening elements. If the temperature is too high, the matrix will penetrate into the dry fiber bundle due to increased flowability.

The alternative manufacturing process requires a combination of materials in which the matrix and resin can be cured / processed at the same temperature. All structural elements of the fiber composite component are then cured in a single process step at a defined temperature (125 ° C with a rubber matrix: AA6CFZ from Kraiburg and epoxy resin HexPly913 from Hexcel) and an atmospheric pressure in an autoclave (4 - 7 bar) .

• 1 schematisch eine Flügelvorderkante eines erfindungsgemäßen Flügels und den zugehörigen Aufbau der Flügelhaut und
• 2 einen alternativen Aufbau einer Flügelhaut für einen erfindungsgemäßen Flügel,
• 3 die Flügelvorderkante im verformten und im unverformten Zustand und
• 4 eine Vorrichtung zum Herstellen (nach dem ersten Verfahren) eines erfindungsgemäßen Faserverbundbauteils, beispielsweise eines erfindungsgemäßen Flügels.
• 5 einen erfindungsgemäßen Flügel mit Hochauftriebsklappe und
• 6 zeigt schematisch ein Flugzeug mit einem erfindungsgemäßen Flügel.

The invention is explained in more detail below with reference to the attached figures. It shows

• 1 schematically a wing leading edge of a wing according to the invention and the associated structure of the wing skin and
• 2nd an alternative structure of a wing skin for a wing according to the invention,
• 3rd the leading edge of the wing in the deformed and in the undeformed state and
• 4th a device for producing (according to the first method) a fiber composite component according to the invention, for example a wing according to the invention.
• 5 a wing according to the invention with high lift flap and
• 6 shows schematically an aircraft with a wing according to the invention.

1 shows a leading edge of the wing 10th of a schematically drawn variable wing 12th . The leading edge of the wing 10th includes a wing skin 14 and a schematically drawn contour changing device 16 . The contour changing device 16 comprises several transmitter elements 18.1 , 18.2 , 18.3 , April 18 that with the wing skin 14 are connected and in the span direction S run.

The transmitter elements 18th (Reference symbols without a counting suffix refer to all corresponding objects) are above joints and transmission rods 20th with a swivel rod 22 connected that by means of a schematically shown swivel drive 24th can be pivoted. A detailed description of how a contour changing device can be constructed can be found in the patent DE 29 07 912 A1 by Dornier.

In the lower part of 1 a section of the wing skin is shown. The wing skin 14 includes a matrix 26 from a matrix material, in the present case from the rubber compound AA6CFZ from Kraiburg. The wing skin 14 also includes a large number of cross-stiffening bodies that are disjoint from one another 28 , of which the cross stiffeners 28.1 , 28.2 , 28.3 and April 28 are drawn. The cross stiffeners 28 extend along the span direction S .

The wing skin 14 has a longitudinal stiffening element 30th in the form of a flat laminate. Several fibers are schematic 32.1 , 32.2 , ... are drawn in, which all run parallel to each other and are perpendicular to the cross-bracing members 28 extend. A neutral fiber is also shown N . In the drawing at the bottom is the cutout of the wing skin 14 after a contour change in a longitudinal bending plane E shown.

The fibers 32 of the longitudinal stiffening element 30th are in a longitudinal stiffening element matrix 34 embedded, which consists of a thermoset. In the present case, the longitudinal stiffening element matrix 34 formed by an epoxy resin. The fibers 32 are glass fibers. The longitudinal stiffening element 30th has a thickness D of approximately 0.8 mm.

The cross stiffeners 28 each comprise a fiber bundle 36 , for which the English term roving could also be used. For example, the cross stiffener has April 28 a bundle of fibers 36.4 that in a cross stiffener matrix 38.4 is embedded. The fiber bundle 36 are made of carbon fibers, the cross-bracing matrix 38 made of epoxy resin.

In 3rd is the contour of the leading edge of the wing 10th in a coordinate system with the dimensionless wing depth c and the standardized wing thickness t applied once in a state with a small lift coefficient (solid line) and once in a state with an increased lift coefficient (dashed line). By actuating the part-turn actuator 24th the contour of the leading edge of the wing changes 10th how this is shown schematically in 3rd is drawn. The matrix deforms 26 . The area of the matrix 26 who in 1 above the longitudinal stiffening element 30th is arranged, is stretched, for example, the area below it is compressed.

4th shows how to make such a contour-changing wing 12th or a fiber composite component according to the invention initially fiber bundles 36 with a laying head 40 placed on a mat made of thermoplastic elastomer. Another elastomer mat is then placed on top and the laying head is used 40 the fibers 32 for the longitudinal stiffening element 30th placed perpendicular to the fiber bundles placed first. This layer is covered again with a mat made of thermoplastic elastomer. Another layer of fiber bundles follows, which run parallel to the fiber bundles laid first. These are also covered with a mat made of thermoplastic elastomer.

In the next step, this structure is exposed to a vacuum of 0.7 bar and a temperature of 172.5 ° C. This melts the thermoplastic elastomer around the dry fiber bundle 36 . Next, a cold curing resin is placed in the spaces between the dry fiber bundles 36 (by applying a pressure difference). The resin and the fibers 32 form the transverse and longitudinal auction elements and, when hardened, connects them to the matrix 26 .

2nd shows an alternative embodiment of the wing skin 14 , in which the longitudinal stiffening element 30th is arranged outside.

5 shows a wing according to the invention 12th with a high lift flap 42 and a blow-out opening 44 , from which a tangential air flow is blown out during operation. The isobars show the resulting pressure gradients. The high-lift flap 42 comprises a high lift flap movement device, not shown, by means of which the inclination of the flap to the rest of the wing 12th can be changed.

6 shows an airplane 46 with three engines 48.1 , 48.2 , 48.3 of which an adjustable flow of bleed air through an unillustrated channel to the exhaust openings 44 the wing 12.1 , 12.2 , 12.3 , 12.4 is directed.

Reference list

10th

Leading edge

12th

wing

14

Wing skin

16

Contour changing device

18th

Transmitter element

20th

Transmission rod

22

Swivel rod

24th

Swivel drive

26

matrix

28

Cross stiffening body

30th

Longitudinal stiffening element

32

fiber

34

Longitudinal stiffening element matrix

36

Fiber bundle

38

Cross stiffener matrix

40

Laying head [without reference number in 4th ]

42

High-lift flap

44

Discharge opening

46

plane

48

Engine

S

Span direction

D

thickness

A

Flow direction

E

Longitudinal bending plane

N

neutral fiber

c

Wing depth

t

Wing thickness

Claims (9)

Contour-changing wing (12), which extends in a span direction (S) and (i) a wing skin (14) and (ii) a contour changing device (16), by means of which a lift coefficient of the wing (12) by changing a contour of the wing skin ( 14) can be changed on a leading edge (10) of the wing, characterized in that (iii) the wing skin (14) (a) a matrix (26) made of a matrix material, (b) at least one transverse stiffening body (28) which extends along extends in the span direction (S), and (c) has at least one longitudinal stiffening element (30) which extends along an inflow direction (A) of inflowing air during operation of the wing (12), (d) the matrix (26) being a has less rigidity than the longitudinal stiffening element (30), so that when the contour of the wing skin (14) changes in a longitudinal bending plane (E) that is perpendicular to the transverse stiffening bodies (28), a neutral fiber (N) in the longitudinal stiffener tion element (30). Contour-variable wing (12) after Claim 1 , characterized in that the longitudinal stiffening element (30) forms a curved surface which extends along the leading edge of the wing (10). Contour-variable wing (12) according to one of the preceding claims, characterized in that the longitudinal stiffening element (30) has a thickness (D) of at most 2 millimeters, in particular at most 1 millimeter. Contour-variable wing (12) according to one of the preceding claims, characterized in that the at least one cross-stiffening body (28) is constructed from a fiber composite material and has a multiplicity of fibers (32) running in parallel in a cross-stiffening body matrix (38). Contour-variable wing (12) according to one of the preceding claims, characterized by at least one second transverse reinforcement body (28.4) which extends along the span direction (S), the longitudinal reinforcement element (30) between the first transverse reinforcement body (28.1) and the second transverse reinforcement body (28.4 ) is arranged. Method for producing a contour-changing wing (12), comprising the steps: (i) Surrounding fiber bundles (36), which consist of a plurality of fibers (32), with a matrix (26) made of elastomer, so that covered fiber bundles (36) are formed, and (ii) subsequently consolidating the dry fibers (32) of the coated fiber bundles (36) with a resin, so that the resin connects the fibers (32) to one another and forms a connection with the matrix (26). Procedure according to Claim 6 characterized in that consolidating the fibers (32) comprises the steps of: applying a pressure difference between a first end of the covered fiber bundle (36) and a second end of the covered fiber bundle (36) opposite the first end, and - Sucking through and / or pushing the resin through the fiber bundle (36). Method for producing a contour-changing wing (12), comprising the steps: - Surrounding a plurality of fiber bundles (36), which are pre-impregnated with a resin, with a matrix material made of an elastomer and - Simultaneously heating the fiber bundles (36), the resin and the matrix material, so that the fiber bundle (36) is melted by the matrix and the pre-impregnated fiber bundles (36) form a connection with the matrix (elastomer) (26) by melting resin. Fiber composite component, characterized by (a) a matrix (26) made of a matrix material, (b) a plurality of transverse stiffening bodies (28) which - each have a fiber bundle (36) which is bound by a resin, - are embedded in the matrix (26), - each have a longitudinal extension and - are arranged parallel to one another, and (c) at least one longitudinal stiffening element (30) which - has a multiplicity of fibers (32) which are bound by a resin and which run essentially parallel to one another, and - runs at an angle of at least 75 ° to the transverse stiffening bodies (28), (d) the matrix (26) has a lower rigidity than the longitudinal stiffening element (30), so that when the contour of the fiber composite component changes in a longitudinal bending plane (E), which is perpendicular to the transverse stiffening bodies (28), a neutral fiber (N) runs in the longitudinal stiffening element (30) and ( e) wherein a transverse bending stiffness with respect to a bend in a transverse bending plane bet at least eight times a longitudinal bending stiffness with respect to a bend in the longitudinal bending plane (E) struggles.

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