EP2989003A1

EP2989003A1 - Stiffening element run-out

Info

Publication number: EP2989003A1
Application number: EP13882746.4A
Authority: EP
Inventors: Stefan Gustavsson
Original assignee: Saab AB
Current assignee: Saab AB
Priority date: 2013-04-25
Filing date: 2013-04-25
Publication date: 2016-03-02
Also published as: EP2989003A4; WO2014175798A1

Abstract

The invention regards a torsion-box type airfoil composite structure comprising a skin component (6, 206) and at least a first and a second stiffening element (8, 8', 208', 8'', 208'') arranged adjacent to each other longitudinally along the skin component (6, 206), wherein the at least first and second stiffening element (8, 8', 208', 8'', 208'') each comprise a flange member (14', 214', 14'', 214'') arranged in parallel with and bonded to the skin component (6, 206), wherein the second stiffening element (8, 8'', 208'') comprises a run-out area (10, 210). The flange member (14', 214') of the first stiffening element (8, 8', 208') and the flange member (14'', 214'') of the second stiffening element (8, 8'', 208'') are joined at the run-out area (10, 210).

Description

Stiffening element run-out

TECHNICAL FIELD The present invention relates to a torsion-box type airfoil composite structure according to the preamble of claim 1 and a method according to claim 13 for manufacturing a torsion-box type composite structure. The invention also regards a data medium storing program comprising a program code, which program when run on a computer executes the method according to the invention. The invention also regards a production line per se adapted to make use of the data medium storing program for executing the method.

The invention relates to the aircraft industry and to aircraft service engineering. The invention is not limited thereto, but could also be related to activities regarding main- tenance of commercial aircraft as well.

BACKGROUND ART

Torsions boxes of airfoils consist of several structural elements such as upper and lower skins, stiffening elements (stringers), spars and ribs. The skin is stiffened with elongated stiffening elements, such as stringers, arranged longitudinally along the skin, which improves the strength and the buckling behaviour of the skin. Due to the shape of the airfoil and conflicts with other structural elements some stiffening elements have to be terminated before others and the stiffening elements therefore have different lengths. At the termination of a stiffening element, the stiffening element run- out, the load must be transferred from the stiffening element to the skin. The redistribution of loads at the run-out causes a local stress concentration which affects the bonding between the stiffening element and the skin. The bonding strength at the runout is therefore crucial and stiffening elements are often designed to comprise a runout region to transfer the load in such a way as to minimise the stress. With the use of composite materials, such as carbon fibre-reinforced plastic, the structural performance at the run-out is of particular interest. Specifically, in composite structures where stiffening elements are co-cured with the skin, the run-outs are prone to be subject to cracking due to the natural offset between the stiffening element and the skin. Addi- tional elements for load transfer or for strengthening the bond between the stiffening element and the skin are therefore often used at the run-out region.

US 2010/0127122 A1 discloses a composite structure which may form, for example, the skin of an aircraft wing. The structure comprises a panel and a series of stringers (stiffening elements) bonded to the surface of the panel. At a stringer run-out a pad protrudes downwardly from the base of the stringer and extends beyond the ends of the web and flanges. The pad is embedded in a recess in the panel and fasteners such as bolts may be employed to fix the pad into the recess. The flange has a ta- pered upper face at the run out to ensure a smooth load transfer from the stringer via the pad to the panel.

Furthermore, US 2012/0234978 A1 discloses a composite structure comprising a device for transferring load of a stringer (stiffening element) to the skin in a stringer run- out zone of an aircraft. The device comprises two metal brackets to be joined the each side of the stringer web and feet having a first section to be joined to each side of the stringer foot and a second section to be joined to the skin. The device is joined to the skin by using fasteners, such as bolts. By using for example bolts or rivets which penetrate the skin, the outer surface (aerodynamic surface) of the skin (wing skin, fuselage skin, stabilizer skin, fin skin, nacelle skin etc.) will exhibit rivet or bolt heads being exposed to the airflow. The heads will disturb the natural flow of air flowing over the outer surface of the skin when the airfoil structure is used and turbulence will occur. Turbulence increases the aerodynamic drag. Also, by using an extra component for transferring the load, additional weight is added to the structure. An aircraft comprising an airfoil with increased aerodynamic drag and increased weight will have increased fuel consumption.

The object of the present invention is to provide a laminar airfoil composite structure with high strength stiffening element run-outs.

Another object of the invention is to provide an airfoil composite structure which is compact and which is of lower weight relative prior art structures but still provide high strength. Another object of the invention is to provide an airfoil composite structure comprising integrated stiffening element run-outs. Another object of the invention is to provide an airfoil composite structure with reduced weight compared to known technology.

Another object is to provide an aircraft, which has low fuel consumption, and therefore can be regarded as a green technology.

Another object of the invention is to provide an airfoil composite structure which can be manufactured in an automated production line.

Another object is to provide a method for producing a laminar (aerodynamic smooth outer surface of the structure) torsion-box type airfoil composite structure.

Another object is to provide a data medium storing program that, when it is executed on a computer together with a production line data system or included in it, enables an automatic or semi-automatic execution of the various stages of the aforementioned method.

The foregoing and other objects and advantages of the present invention will be apparent to those skilled in the art, in view of the following detailed description, taken in conjunction with the appended claims and the accompanying drawings.

SUMMARY OF THE INVENTION

The objects of the invention have been achieved by the composite structure defined in the introduction and characterized by the features of the characterizing part of claim 1.

By joining the flange members of the first and the second elongated stiffening elements at the run-out area of the second stiffening element, the second stiffening element run-out is joined both to the adjacent first stiffening element and to the skin component. Thus is achieved a torsion-box type airfoil composite structure comprising high strength integrated stiffening element run-outs.

Preferably, the at least first and second stiffening elements are bonded to each other and the skin component by co-curing. Thus is provided a laminar airfoil composite structure which is compact and integrated and which comprises integrated stiffening element run-outs. Alternatively, an adhesive is also used between the stiffening elements and the skin component to strengthen the bond and to eliminate any gaps and differences between the surfaces.

In such way no fasteners, such as bolts or rivets, are needed to strengthen the runouts which entail an airfoil composite structure comprising a smooth outer surface. The smooth outer surface causes a natural laminar airflow over the airfoil composite structure which reduces the aerodynamic drag. The airfoil composite structure accord- ing to the invention does not require any additional fittings at the stiffening element run-out and the structure can thereby sustain a relatively low weight compared to known technology. Reduced aerodynamic drag and reduced weight results in lower fuel consumption. The flange members of the first and the second stiffening element are preferably joined at the run-out area by a scarf joint. The scarf joint can suitably be configured as correspondingly stepped or bevelled faces of the flange members of the first and the second stiffening elements joined to each other. The stepped faces can each comprise one or a plurality of steps. Thus is achieved a high strength integrated stiffening element run-out. Alternatively, the flange members of the first and the second stiffening elements are joined by a butt joint at the run-out area.

The bevelled or stepped joining faces in the scarf joint are preferably configured with a ratio (height/length) between 1 :10 to 1 :40.

Suitably, the flange member of the second stiffening element is extended in the longitudinal direction beyond the web of the second stiffening element. The extended part of the flange member preferably comprises the stepped or bevelled face forming a part of the scarf joint. The stiffening elements are preferably arranged substantially in parallel to each other.

Preferably, at the run-out area where the second stiffening element terminates, the flange member of the first stiffening element extends substantially transversally to the elongation of the first stiffening element, parallel to the skin component, providing an extended part. The extended part of the flange member of the first stiffening element thus extends into the run-out area, in front of the second stiffening element seen in the longitudinal direction, across the width of the flange member of the second stiffening element, substantially at right angle to the length of the second stiffening element. The edge of the extended part facing the second stiffening element is joined to the flange member of the second stiffening element along at least a portion of the width of the flange member of the second stiffening element (at the run-out area). By joining the flange member of the first stiffening element with the flange member of the second stiffening element, by means of a joint extending transversely to the longitudinal direction of the first and the second stiffening elements, loads acting upon the second stiffening element in the longitudinal direction can be absorbed efficiently in the run-out area. Alternatively, the flange member of the first stiffening element and the flange member of the second stiffening element are configured such that they are joined, by means of a joint extending at an angle to the longitudinal direction (the length) of the first and the second stiffening elements. The angle is suitably greater than 0 degrees and smaller than 180 degrees, preferably between 40-55 degrees or 130-145 degrees.

The flange member of the first stiffening element and the flange member of the second stiffening element are preferably arranged in abutment along the length of the second stiffening element and are preferably longitudinally joined by a butt joint. Alternatively the flange member of the first stiffening element and the flange member of the second stiffening element are longitudinally joined by a scarf joint.

In this way the flange members of the at least first and the second stiffening elements form a composite layer, which is covering the skin component. The layer formed by the flange members is bonded to the skin component and is thus integrated with the skin component. The thickness of the layer formed by the flange members supplement and adds to the thickness of the skin component and the skin component can therefore be provided thinner than in known technology. Thus is achieved a laminar airfoil composite structure with reduced weight compared to known technology.

Suitably, the first and the second stiffening elements are configured with a T cross- section such that the flange member of each stiffening element is arranged in parallel with and bonded to the skin component, and such that each stiffening element comprises a web, extending at substantially right angle to the skin component.

Preferably, the lower face the flange member of the second stiffening element is stepped or bevelled at the run-out area while the upper face of the edge of the extended part of flange member of the first stiffening element facing the second stiffening element is correspondingly stepped or bevelled, such that a scarf joint is formed when joining the stepped or bevelled faces. Alternatively the lower face of the edge of the extended part of the flange member of the first stiffening element facing the second stiffening element and the upper face of the flange member of the second stiffening element are correspondingly stepped or bevelled. In such way, a laminar airfoil composite structure is provided, which comprises a high strength integrated stiffening element run-out.

The T-shaped stiffening elements are preferably formed by two L-shaped profiles placed back-to-back with a blade arranged in between. The blade increases the thick- ness of the webs of the stiffening elements and thus increases the stability of the structure. Alternatively, the stiffening elements are formed by one profile shaped as a T.

The at least first and second stiffening elements are alternatively configured with omega or Z, C, H or U cross-sections or other profiles.

Preferably, the web of the second stiffening element has a tapered portion at the runout area to facilitate load transfer from the stiffening element to the skin component at the run-out area. The web is tapered such that the height of the web is reduced to- wards the terminating end (run-out) of the stiffening element. The ratio (height of the web/length of the tapered portion) of the tapered portion is 1 :2— 1 :20, preferably 1 :3 - 1 : 10. Preferably, the airfoil composite structure comprises fastening elements (e.g. rib feet) for attachment of the composite structure to other structural components (e.g. wing ribs). The fastening elements are suitably configured having a T-shape, such that each fastening element comprises a web and a flange member. The fastening elements are preferably arranged such that the plane of the flange member is parallel with the plane of the skin component and the web extends at substantially right angle to the skin component.

The fastening elements are preferably arranged between the stiffening elements, substantially at right angles to the elongation (length) of the stiffening elements. A fasten- ing element is suitably arranged within the run-out area of the second stiffening element for additionally strengthening the run-out of the stiffening element.

Preferably, the airfoil composite structure is configured as an upper face of an aircraft wing. The stiffening elements are preferably stringers and the fastening elements are preferably rib feet for attachment of the structure to ribs. The run-out of the second stringer (stiffening element) is preferably arranged adjacent a rib foot and also in proximity to a spar member for improved structural stability. The airfoil composite structure is preferably configured to comprise a series of stiffening elements arranged substantially parallel to each other wherein at least two stringers comprise run-out areas.

The stiffening elements, fastening elements and the skin component are preferably formed by lay-ups comprising stacks of pre-preg plies. Preferably, each layer of ply of the skin component lay-up, the stiffening element lay-ups and the fastening element lay-ups is provided by an automated tape laying (ATL) machine.

The lay-up material being used to form the stiffening elements, fastening elements and the skin component may be of any suitable resin pre-impregnated fibre material. The lay-ups may be based on unidirectional pre-impregnated fiber plies, the fibers being of woven carbon fiber pre-preg fabrics, Kevlar, spectra pre-preg tapes and fabrics etc. Alternatively, the material used can be dry woven fabric, in a second step injected or impregnated with resin.

Suitably, the material used can be dry woven fabric interleaved with resin film. This has also been achieved by a method as claimed by the steps of claim 13. The method preferably comprises the steps of providing a forming tool, providing a skin component lay-up onto the forming tool, positioning a stiffening element lay-up on the skin component lay-up forming an integrated lay-up with integrated run-outs and co- curing the integrated lay-up.

The method preferably also comprises the step of positioning a fastener element lay- up on the skin component prior to the co-curing.

Preferably, the cure tool comprises an autoclave apparatus.

The step of positioning the stiffening element lay-up on the skin component lay-up is preferably performed after forming said stiffener lay-up over a forming tool.

The stiffening elements are preferably formed using an elongated forming tool suitably made from aluminium or any other hard tool material. The forming tool is preferably configured as a modulate box tool, comprising a series of box tools each comprising a forming surface.

The stepped faces of the flange members of the stiffening elements, forming a scarf joint, are preferably achieved by varying the length of the pre-preg plies in the lay-up.

Preferably, a supporting wedge is used in combination with the forming tool when forming a stiffening element comprising a stepped face. The supporting wedge is suitably stepped correspondingly to the desired stepped face. Alternatively, the form- ing surface of the forming tool is correspondingly stepped in order to support the stepped stiffening element lay-up.

The stiffening element lay-up and the fastening element lay-up are preferably pro- vided and formed on the same forming tool and are thus positioned on the skin component lay-up at the same time.

The step of positioning the stiffening element lay-up onto the skin component lay-up is preferably performed such that the flange member of a first stiffening element and the flange member of a second stiffening element are joined within the area of the second stiffening element run-out.

The step of positioning the stiffening element lay-up onto the skin component lay-up is preferably performed such that the flange member of the first stiffening element and the flange member of the second stiffening element are longitudinally joined.

The method preferably comprises the step of providing an adhesive between the stiffening element lay-up and the skin component lay-up before positioning the stiffening element lay-up on the skin component lay-up.

This is also solved by an aircraft wing comprising a torsion-box type airfoil composite structure as defined in the introduction, characterized by the features of any of claims 1-11 , wherein the airfoil composite structure is attached to a main sub-structure. In such way is achieved an aircraft wing, which comprises a smooth outer surface at the same time as the bonding of the stiffening element run-outs to the skin component is strengthened.

This has also been achieved by a production line for the production of a torsion-box type airfoil composite structure according to claim 16.

The production line preferably comprises means for positioning fastening element lay- ups on the skin component lay-up prior to curing. In such way is achieved a production line which is suitable for automation, which production line provides a torsion-box type airfoil composite structure cost-effectively in such way that the outer surface of the airfoil composite structure is smooth and still achieves a rigid stiffening element run-out. This is cost-effective and time saving. Such production line will permit a high production rate of the manufacture of the torsion-box type airfoil composite structure.

The means for providing the forming tool can be an automatic machining apparatus adapted to machine the forming tool. The means for providing a skin component lay- up and/or the stiffening element lay-up and/or the fastening element lay-up onto the forming tool may be an ATL apparatus operated semi-automatic or automatic. The means for positioning the stiffening element lay-up and the fastening element lay-up on the skin component lay-up forming an integrated lay-up can be a robot cooperating with an automatic forming tool for forming stiffening element lay-ups one by one and fastening element lay-ups one by one and thereafter automatically or semi- automatically arranging the lay-ups on a box tool modular set up. The means for co- curing the integrated lay-up may be an autoclave into which the integrated lay-up automatically or semi-automatically is inserted. Thereafter the integrated lay-up is demoulded, wherein the box tools are removed from the lay-up. This can be performed by a robot arm connected to a gripping device gripping the box tools.

This is solved also by a data medium storing program according to claim 18 for establishing in a production line according to claim 16 or 17 an automatic or semi-automatic manufacture of a torsion-box type airfoil composite structure.

This is also solved by a data medium storing program product according to claim 19 comprising a program code stored on a medium, which is readable on a computer, for performing the method steps according to any of claims 13 to 15, when a data medium storing program according to claim 18 is run on a control unit.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention will now be described by way of examples with references to the accompanying schematic drawings, of which: Fig. 1 illustrates an aircraft comprising a torsion-box type airfoil composite structure according to the invention;

Fig. 2 illustrates a cross-sectional view of an aircraft wing comprising the torsion-box type airfoil composite structure according to the invention;

Fig. 3 illustrates a side view of a torsion-box type airfoil composite structure according to one embodiment of the invention;

Fig. 4a illustrates a side view of a run-out area of a torsion-box type airfoil composite structure according to another embodiment of the invention;

Fig. 4b illustrates a top view of a run-out area of a torsion-box type airfoil composite structure according to one aspect of the invention;

Fig 5a-5b illustrate details of a scarf joint according to the invention;

Fig. 6 illustrates a side view of a torsion-box type airfoil composite structure according to a further embodiment of the invention;

Fig. 7 illustrates a side view of a torsion-box type airfoil composite structure according to an additional embodiment of the invention;

Fig. 8a-8c illustrates flowcharts for methods for manufacturing a torsion-box type airfoil composite structure according to one aspect of the invention;

Fig. 9a-9d illustrates production tools for producing details of a torsion-box type airfoil composite structure according one aspect of the invention;

Fig. 10a-10e illustrates a production line for producing a torsion-box type airfoil composite structure according to one aspect of the invention; and

Fig. 10f illustrates a device according to one aspect of the invention. DETAILED DESCRIPTION

Hereinafter, embodiments of the present invention will be described in detail with reference to the accompanying drawings, wherein for the sake of clarity and understanding of the invention some details of no importance are deleted.

Fig. 1 shows an aircraft 1 comprising a torsion-box type airfoil composite structure 2 according to one aspect of the invention. The torsion-box type airfoil composite structure 2 forms a part of an aircraft 1 wing 4 to stiffen the wing 4 and to provide a laminar wing 4 with a smooth outer surface which reduces the aerodynamic drag. Fig. 2 illustrates a part of an aircraft 1 wing 4 comprising the torsion-box type airfoil composite structure 2 according to one aspect of the invention. The composite structure 2 comprises a skin component 6 and a series of elongated stringers 8 (stiffening elements) bonded to the skin component 6. The stringers 8 are arranged substantially parallel to each other and exhibit an extension in span wise direction of the wing 4. Due to the shape of the wing 4 the stringers 8 will have different lengths and some stringers 8 therefore comprise run-out areas 10. At the run-out area 10 the load acting upon the stringer 8 must be transferred to the skin component 6. The run-out area 10 is therefore subject to a local stress concentration, which affects the bonding between the stringer 8 end and the skin component 6. The run-out area 10 should be designed such as to minimise the stress. At the run-out area 10, adjacent stringers 8 are therefore joined to each other substantially transversally to the longitudinally direction (chord wise) by scarf joints as described in Fig. 3 - Fig. 7. The torsion-box type airfoil composite structure 2 further comprises a spar member 9 wherein the stringer 8 runouts are arranged adjacent the spar member 9 to improve the structural stability of the composite structure 2 and to strengthen the stringer 8 run-out area 10.

Fig. 3 shows a part of a torsion-box type airfoil composite structure 2 according to one aspect of the invention. The structure 2 comprises a skin component 6 and at least a first and a second elongated stiffening element 8', 8" (stringer) stiffening the skin component 6. The at least first and second stiffening elements 8', 8" are arranged adjacent each other and runs substantially parallel with each other longitudinally (span wise) along the skin component 6. The second stiffening element 8" is configured such that it is shorter than the first stiffening element 8' and comprises a run-out area 10. The first stiffening element 8' and the second stiffening element 8" are preferably each embodied having T-shapes, such that a web 12', 12" is extending at right angle to the skin component 6 and a flange member 14', 14" is parallel with and bonded to the skin component 6. The flange members 14', 14" each comprise a first flange 16' (shown in Fig. 4b), 16" and a second flange 18', 18". The T-shaped stiffening elements 8', 8" are preferably formed by two L-shaped profiles placed back-to-back with a blade 20', 20" arranged in between. The blade 20', 20" increases the thickness of the webs 12', 12" of the first and the second stiffening elements 8', 8" and thus increases the stability of the structure 2. The second flange 18' of the first stiffening element 8' and the first flange 16" of the second stiffening element 8" are arranged in abutment along the length of the second stiffening element 8" and are longitudinally joined by a butt joint. The web 12" of the second stiffening element 8" comprises a tapered portion 22 within the run-out area 10, for transferring loads from the second stiffening element 8" to the skin component 6. The web 12" is tapered such that the height of the web 12" is reduced towards the run-out. Where the second stiffening element 8" run out, the second flange 18' of the first stiffening element 8' extends transversely to the longitudinal direction, parallel to the skin component 6, such that it extends in front of the run-out area 10 seen in the longitudinal direction. The second flange 18' of the first stiffening element 8' thus provides an extended part 24, which extends across the width of the second stiffening element 8", substantially at right angle to the length (elongation) of the second stiffening element 8" and is joined to the first and the second flanges 16", 18" of the second stiffening element 8". The edge 26 of the extended part 24, facing the second stiffening element 8", is joined to the first and the second flanges 16", 18" of the second stiffening element 8" by a scarf joint extending transversely to the longitudinal direction (chord wise). In this way the flange members 14', 14" of the first and the second stiffening element 8', 8" form a composite layer covering the skin component 6. The flange members 14', 14" are thus rigidly bonded to each other and to the skin component 6 by co-curing and thus become an integrated part of the skin component 6. The integrated stiffening elements 8', 8" and stiffening element run-out comprising the chord wise scarf joint entails a high strength run-out without the need of additional fasteners protruding through the skin component 6 and thereby provides a laminar airfoil composite structure. Fig. 4a shows an embodiment of a run-out area 10 of the airfoil composite structure 2 according to Fig. 3 where the scarf joint is formed by stepped faces. The first and the second flanges 16", 18" of the second stiffening element 8" extend longitudinally beyond the web 12" of the second stiffening element 8". The upper face of the extended part 24 of the second flange 18' of the first stiffening element 8' comprises a stepped portion 28' at the edge 26 facing the second stiffening element 8". The lower face 28" of the extended section of the first and the second flanges 16", 18" of the second stiffening element 8" are correspondingly stepped such that it matches the stepped portion 28' of the extended part 24 of the first stiffening element 8'. The stepped chord wise scarf joint provides a high strength run-out when the stiffening elements 8', 8" are co-cured with the skin component 6.

Fig. 4b shows a top view of a run-out area 10 of the airfoil composite structure 2 ac- cording to Fig. 3 according to one aspect of the invention. The second flange 18' of the first stiffening element 8' and the first and the second flanges 16", 18" of the second stiffening element 8" are configured such that they are joined by means of a scarf joint extending with an angle a to the longitudinal direction (the direction of the elongation of the stiffening elements). The angle a is greater than 0 degrees and smaller than 180 degrees. Preferably, the angle a is 45 degrees.

Fig. 5a and 5b illustrate details of two embodiments of a scarf joint joining the first stiffening element 8' and the second stiffening element 8" at the run-out area 10 according to one aspect of the invention. Fig. 5a shows the scarf joint formed by stepped faces 28', 28" of the flange members 14', 14". The first and the second stiffening elements 8', 8" are formed by lay-ups 29 comprising stacks of pre-preg plies 30. The stepped faces 28', 28" are formed by sequentially laying pre-preg plies 30 with different lengths on top of each other when forming the stiffening elements 8', 8". Fig. 5b shows an alternative embodiment where the scarf joint is formed by bevelled faces 32', 32".

Fig. 6 illustrates a side view of an alternative embodiment of the torsion-box type airfoil composite structure 2 described in Fig. 3. Fastening elements 34 (rib feet) are arranged between the stiffening elements 8', 8" (stringers) substantially transversally to the length of the stiffening elements 8', 8". The fastening elements 34 are preferably T-shaped and each comprises a web 36 and a flange member 38. The flange members 38 of the fastening elements 34 are arranged parallel with the skin component 6 and on top of the flange members 14', 14" of the stiffening elements 8', 8". The web 36 of the fastening element 34 extends substantially at right angles to the skin component 6. The run-out area 10 of the second stiffening element 8" is arranged adjacent a fastening element 34 such that the flange member 38 of the fastening element 34 covers a part of the first flange 16" of the second stiffening element 8" within the run-out area 10. Thus, the fastening element 34 improves the structural stability at the run-out area 10. Fig. 7 shows a side view of a torsion-box type airfoil composite structure 2 according to another embodiment of the invention. The structure 2 comprises a skin component 206 and at least a first and a second elongated stiffening element 208', 208" each configured with a Z cross-section. The stiffening elements 208', 208" each comprises a web 212'. 212", extending at substantially right angle to the skin component 206 and a flange member 214', 214" arranged in parallel with and bonded to the skin component 206. The second stiffening element 208" comprises a run-out area 210. The flange member 214' of the first stiffening element 208' and the flange member 214" of the second stiffening element 208" are arranged longitudinally adjacent each other. Preferably the flange member 214' of the first stiffening element 208' and the flange member 214" of the second stiffening element 208" are arranged in abutment and joined by a butt joint. Where the second stiffening element 208" is terminated (run-out), the flange member 214' of the first stiffening element 208' extends trans- versely to the longitudinal direction, in front of the run-out area 210. The flange member 214' of the first stiffening element 208' thus provides an extended part 224, which extends across the width of the flange member 214" of the second stiffening element 208" and substantially at right angle to the length of the second stiffening element 208". The edge 226 of the extended part 224 facing the second stiffening element 208" is joined with the flange member 214" of the second stiffening element 208" by a scarf joint transversely to the longitudinal direction. The flange members 214', 214" are bonded to each other and to the skin component 6 by co-curing and the stiffening elements 208', 208" and the run-out thus become an integrated part of the skin component 6.

Fig. 8a illustrates a flowchart for a method for producing the torsion-box type airfoil composite structure according to one aspect of the invention. The method starts in a Step 101. In Step 102 is provided a method for manufacture of a torsion-box type airfoil composite structure comprising a skin component and at least a first and a second stiffening element placed adjacent to each other longitudinally along the skin component wherein the at least first and second stiffening element each comprise a flange member arranged in parallel with and bonded to the skin component wherein the second stiffening element comprises a run-out area. Step 102 thus comprises the steps providing forming tools 40, 46, 60, 61 (shown in Fig. 9a - 9d and Fig. 10a - 10e), pro- viding a skin component lay-up 62 (Fig. 10a - 10e) onto the forming tool 60, 61 , positioning a stiffening element lay-up 29 (Fig. 5a - 5b) on the skin component lay-up 62 forming an integrated lay-up and co-curing the integrated lay-up. In Step 103 the method is fulfilled and stopped.

Fig. 8b illustrates a flowchart for a method for producing the torsion-box type airfoil composite structure 2 according to one aspect of the invention. The torsion-box type airfoil composite structure 2 comprising a skin component 6 and at least a first and a second stiffening element 8, 8', 8" placed adjacent to each other longitudinally along the skin component 6 wherein the at least first and second stiffening elements 8, 8', 8" each comprise a flange member 14', 14" arranged in parallel with and bonded to the skin component 6 wherein the second stiffening element 8, 8" comprises a run-out area 10 and wherein the flange member 14' of the first stiffening element 8, 8' and the flange member 14" of the second stiffening element 8, 8" are joined at the run-out area 10. Step 201 corresponds to a starting (start-up) of a production line (described in Fig. 10a-10e). Step 202 defines the providing of a first forming tool 60 (male) and a correspondingly shaped second forming tool 61 (female). In Step 203 a skin component lay-up 62 is provided on the first forming tool 60 by the use of an automated tape laying (ATL) machine 64 (Fig. 10a). The skin component lay-up 62 comprises resin pre-impregnated fiber plies (not shown) not yet being cured. The plies are laid with different directions for achieving a proper strength of the finished skin component 6. The skin component lay-up 62 is thus shaped after the first forming tool 60 and will, when completed, form the outer shape of the torsion-box type airfoil composite structure 2. The first forming tool 60 comprising the skin component lay-up 62 is then turned upside down and positioned in the second forming tool 61 in Step 204. When the skin component lay-up 62 is positioned in the second forming tool 61 , the first forming tool 60 is removed. In Step 205 stiffening element lay-ups 29 of different lengths and comprising run-out areas 10 are positioned onto the skin component lay- up 62 in the second forming tool 61. The stiffening element lay-ups 29 are formed on tools 40, 46 (as described in Fig. 9a-9d), which are positioned such that the flange members of the stiffening element lay-ups 29 are parallel with and in contact with the skin component lay-up 62. An adhesive is preferably added between the skin component lay-up 62 and the stiffening element lay-ups 29. The tools 40, 46 comprising the stiffening element lay-ups are positioned such that the flange members of the stiffen- ing element lay-ups 29 are arranged adjacent each other such that the flange members form a layer covering the skin component lay-up 62 and adding to the thickness of the skin component lay-up 62. The tools 40, 46 comprising the stiffening element lay-ups 29 are also positioned such that the run-out of a shorter stiffening element lay- up 29 is arranged in abutment and joined with an extended part of a flange member of an adjacent stiffening element lay-up 29. The stiffening element lay-ups 29 and the skin component lay-up 62 thus form an integrated lay-up on the second forming tool 61. Step 206 defines the sealing of the second forming tool 61 comprising the integrated lay-up and the introduced stiffening element lay-up tools 40, 46 positioning the skin component lay-up 62 and the stiffening element lay-ups 29 within a vacuum bag 80 (Fig. 10d) for evacuation. Step 207 defines the bagging of the second forming tool 61 comprising the integrated lay-up and introduced stiffening element lay-up tools 40, 46 within a vacuum bag 80 for evacuation. In Step 208 the integrated lay-up is co- cured in a curing tool 82 (Fig. 10d) such as an autoclave, such that the stiffening ele- ments 8, 8', 8" and the skin component 6 are bonded to each other. Step 209 corresponds to the removal of the stiffening element tools 40, 46 and an integrated torsion- box type airfoil composite structure 2 is provided. Step 210 defines a stop in the production line where the finished composite structure 2 can be wrapped for protection and transported to a storage facility or can be transported to a work-shop for mounting the torsion-box type airfoil composite structure 2 to another structure forming for example an aircraft wing, fuselage, fin or nacelle.

Fig. 8c illustrates a flowchart for a method for producing a torsion-box type airfoil composite structure 2 according to one aspect of the invention. The torsion-box type airfoil composite structure 2 comprising a skin component 6, 206, at least a first and a second stiffening element 8, 8', 8"placed adjacent to each other longitudinally along the skin component 6 wherein the at least first and second stiffening elements 8, 8', 8" each comprise a flange member 14', 14" arranged in parallel with and bonded to the skin component 6, wherein the second stiffening element 8, 8" comprises a run-out area 10 and wherein the flange member 14' of the first stiffening element 8, 8' and the flange member 14" of the second stiffening element 8, 8" are joined at the run-out area 10. The Step 301 corresponds to a starting (start-up) of a production line (described in Fig. 10a - 10e). Step 302 defines the providing of a first forming tool 60 (male) and a correspondingly shaped second forming tool 61 (female). In Step 303 a skin component lay-up 62 is provided on the first forming tool 60 by the use of an automated tape laying (ATL) machine 64. The skin component lay-up 62 comprises resin pre-impregnated fiber plies (not shown) not yet being cured. The plies are laid with different directions for achiving a proper strength of the finished laminate. The skin component lay-up 62 is thus shaped after the first forming tool 60 and will, when completed, form the outer shape of the torsion-box type airfoil composite structure 2. The first forming tool 60 comprising the skin component lay-up 62 is then turned upside down and positioned in the second forming tool 61 in Step 304. When the skin component lay-up 62 is positioned in the second forming tool 61 , the first forming tool 60 is removed. In Step 305 stiffening element lay-ups 29 of different lengths and comprising run-out areas 10 and fastening element lay-ups 68 (Fig. 9c) are positioned onto the skin component lay-up 62 in the second forming tool 61. The stiffening element lay- ups 29 and the fastening element lay-ups 68 are formed on the same tools 46 (described in Fig. 9c), which are positioned such that the flange members of the stiffening element lay-ups 29 are parallel with and in contact with the skin component lay-up 62. An adhesive is preferably added between the skin component lay-up 62 and the stiffening element lay-ups 29. The tools 46 comprising the stiffening element lay-ups 29 and the fastening element lay-ups 68 are positioned such that the flange members of the stiffening element lay-ups 29 are arranged adjacent each other forming a layer covering the skin component lay-up 62. The tools 46 comprising the stiffening element lay-ups 29 and the fastening element lay-ups 68 are also positioned such that the runout of a shorter stiffening element lay-up 29 is arranged in abutment with an extended part of a flange member of an adjacent stiffening element lay-up 29. The stiffening element lay-ups 29, the fastening element lay-ups 68 and the skin component lay-up 62 thus form an integrated lay-up on the second forming tool. Step 306 defines the sealing of the second forming tool 61 comprising the integrated lay-up and the introduced stiffening element and fastening element lay-up tools 46 positioning the skin component lay-up 62, the stiffening element lay-ups 29 and the fastening element lay- ups 68 within a vacuum bag 80 (Fig. 10d) for evacuation. Step 307 defines the bag- ging of the second forming tool 61 comprising the integrated lay-up and introduced stiffening element and fastening element lay-up tools 46 within a vacuum bag 80 for evacuation. In Step 308 the integrated lay-up is co-cured in a curing tool 82 (Fig. 10d), such as an autoclave, such that the stiffening elements 8, 8', 8", fastening elements 34 and the skin component 6, 206 are bonded to each other. Step 309 corresponds to the removal of the stiffening element and fastening element tools 46 and an integrated torsion-box type airfoil composite structure 2 is provided. Step 310 defines a stop in the production line where the finished composite structure 2 can be wrapped for protection and transported to a storage facility or can be transported to a work-shop for mounting the torsion-box type airfoil composite structure 2 to another structure forming for example an aircraft wing, fuselage, fin or nacelle.

Fig. 9a illustrates a side view of tools 40 for forming stiffening elements 8 and thus used in a method for manufacturing a torsion-box type airfoil composite structure 2 described in Fig. 8a-8c according to one aspect of the invention. The elongated tools 40 are preferably made from aluminium. A stiffening element lay-up 29 is provided on the tool 40 and bended over the longitudinal edges of the tool 40 such that the stiffening element lay-up 29 is provided with a U-shape. By placing two tools 40 longitudinally in abutment, the adjacent bent edges of two stiffening element lay-ups 29, placed in abutment, form a web of a stiffening element 8. The lay-up 29 comprises a plurality of pre-preg plies 30 (shown in Fig. 5a - 5b), each having a fiber orientation tailor made for each application. The tool 40 could be manufactured semi- automatically or automatically by means of an automatic miller machine coupled to a central control unit (not shown).

Fig. 9b illustrates a side view of tools 40 for forming stiffening elements 8, 8', 8", and thus used in a method for manufacturing a torsion-box type airfoil composite structure 2 described in Fig. 8a-8c according to one aspect of the invention. The elongated tools 40 are preferably made from aluminium. Two stiffening element lay-ups 29 are provided onto each forming tool 40, along the length of each tool 40 and the stiffening element lay-ups 29 are bended over the opposite longitudinal edges of the tool 40. This way is provided two mirror L-shaped stiffening element lay-ups 29 on each tool 40. When placing two tools 40 longitudinally in abutment such that the L-shaped lay- ups 29 are arranged back-to-back, T-shaped elongated stiffening element lay-ups 29 are provided. The lay-ups 29 comprise a plurality of pre-preg plies 30 (shown in Fig. 5a - 5b), each having a fiber orientation tailor made for each application. The tool 40 could be manufactured semi-automatically or automatically by means of an automatic miller machine coupled to a central control unit (not shown). Fig. 9c illustrates a side view of tools 46 for forming stiffening elements 8, 8', 8", and fastening elements 34 and thus used in a method for manufacturing a torsion-box type airfoil composite structure 2 described in Fig. 8a-8c according to one aspect of the invention. The tool 46 comprises a plurality of connected box tools 48. The box tools 48 are preferably made from aluminium. A fastening element lay-up 68 is provided and bended over a longitudinal edge of a separate hard box tool 48 forming an L- shaped fastening element lay-up 68. A plurality of box tools 48 comprising L-shaped fastening element lay-ups 68 are then arranged with their longitudinal edges facing each other, such that the L-shaped lay-ups 68 are arranged back-to-back. Thus are T- shaped fastening elements lay-ups 68 provided. The row of box tools 48 forms an elongated tool 46 with a longitudinal direction at right angle to the fastening element lay-ups 68. In a further step, two stiffening element lay-ups 29 are provided along the row of box tools 48, covering at least a portion of the fastener element lay-ups 68. The stiffening element lay-ups 29 are bended over the longitudinal edges of the tool 46 such that L-shaped stiffening element lay-ups 29 are provided along the length of the tool 46. When positioning two rows of box tools 48 (tools 46) in parallel and in abutment, such that the L-shaped stiffening element lay-ups 29 are arranged back-to- back, T-shaped elongated stiffening element lay-ups 29 are provided. Preferably, a blade 20', 20" (shown in Fig. 3) is positioned between two parallel rows of box tools 48 such that the web of the T-shaped stiffening element lay-ups 29 provided comprises the blade 20', 20", which increases the structural stability of the web. The lay- ups 29 comprise a plurality of pre-preg plies 30 (shown in Fig. 5a - 5b), each having a fiber orientation tailor made for each application. The tool 46 could be manufactured semi-automatically or automatically by means of an automatic miller machine, CNC, coupled to a central control unit (not shown).

Fig. 9d illustrates an enlarged side view of a tool 40, 46 for forming stiffening elements 8, 8', 208', 8", 208" and thus used in a method for manufacturing a torsion-box type airfoil composite structure 2 described in Fig. 8a-8c according to one aspect of the invention. The flange members of two adjacent stiffening elements lay-ups 29 are joined by a scarf joint at a stiffening element run-out area 10, 210, as illustrated in Fig. 3, 4, 5a, 5b and 7. The scarf joint can be a stepped scarf joint formed by correspondingly stepped faces 28', 28" of the stiffening elements 8, 8', 208', 8", 208". The stepped faces 28', 28" are preferably formed by stiffening element lay-ups 29 com- prising pre-preg plies 30 of different lengths as shown in Fig. 5a. When forming stiffening elements 8, 8', 208', 8", 208" comprising such stepped faces 28', 28" a tool 40, 46 is used, such as described in Fig. 9a-9c. When forming a stiffening element 8, 8', 208', 8", 208" comprising a stepped lower face the pre-preg ply 30 (shown in Fig. 5a - 5b) provided first on the tool 40, 46 is the longest while the following provided plies 30 are shorter and shorter. When forming a stiffening element 8, 8', 208', 8", 208" comprising a stepped upper face the pre-preg ply 30 provided first on the tool 40, 46 is the shortest followed by longer and longer plies 30. This way, each provided ply 30 will be longer than the previously provided ply 30 and will not be completely supported by the tool 40, 46 or any underlying ply 30 which could cause a bending of the stepped lay- up 29. In order to support the stepped lay-up 29 a supporting correspondingly stepped wedge 50 is placed adjacent the lay-up 29. That way, the extending part of each added ply 30 is supported by a step of the wedge 50 arranged under the plies 30. Fig. 10a-e schematically illustrates method steps and manufacture means of a production line PL for manufacturing the torsion-box type airfoil composite structure 2 according to one aspect of the invention. A control unit 500 is adapted to control the manufacture steps performed by the production line PL. The control unit 500 is manoeuvred by an operator (not shown). The production line PL is arranged to perform the manufacture steps of providing forming tools 60, 61 , 40, 46, providing a skin component lay-up 62 onto the forming tool 60, positioning stiffening element lay-ups 29 and fastening element lay-ups 68 on the skin component lay-up 62 forming an integrated lay-up and co-curing the integrated lay-up. Fig. 10a shows means for providing a skin component lay-up 62 onto a first forming tool 60. This is achieved by use of an automated tape laying (ATL) machine 64 which is semi-automatically or automatically operated for the application of a skin component lay-up 62 onto a forming surface of a first forming tool 60. The skin component lay-up 62 comprises with resin pre- impregnated fiber plies (not shown) not yet being cured. The plies are laid with different directions for achieving a proper strength of the finished laminate. The control unit 500 operates the ATL machine 64. Fig. 10b shows means for turning the first forming tool 60 comprising the skin component lay-up 62 upside down and positioning the skin component lay-up 62 onto a correspondingly shaped second forming tool 61. A set of elongated actuators 66 (such as hydrauls) are operated by the control unit 500 to move the skin component lay-up 62 into position onto a forming surface of the second forming tool 61. The actuators also removes the first forming tool 60 when the skin component lay-up has been correctly positioned on the second forming tool 61. Fig. 10c shows means for positioning stiffening element lay-ups 29 and fastening element lay-ups 68, on the skin component lay-up 62 forming an integrated lay-up. A set of elongated actuators 66 (such as hydrauls) are operated by the control unit 500 to position tools 46 comprising the stiffening element lay-ups 29 and the fastening element lay-ups 68 such that the stiffening element lay-ups 29 and the fastening element lay- ups 68 are arranged onto the skin component lay-up 62 inner side. Fig. 10d shows means for co-curing the integrated lay-up comprising the skin component lay-up 62, the stiffening element lay-ups 29 and the fastening element lay-ups 68. The integrated lay-up is assembled together with sub-spar tools (not shown), blade lay-ups (not shown) of pre-pregs inserted within stiffening element lay-ups 29 and fastening element lay-ups 68, noodles and other suitable components constituting an assembly. The assembly and second forming tool 61 are bagged in a vacuum bag 80, wherein air suction is performed for sucking eventual air from the pre-preg ply interfaces, thus compacting the lay-ups 29, 68, 62 and preparing the assembly for a proper co-curing. After the bagging procedure, the assembly comprising the integrated lay-up is automatically inserted into an autoclave tool 82 for co-curing. The autoclave 82 provides that the integrated lay-up co-cures in hot air temperature and under high over pressure. Thereafter, the second forming tool 61 with the co-cured laminate is removed from the autoclave 82. Fig. 10e shows means for removing the stiffening element and fastening element lay-up tool 46 after the co-curing. This is preferably achieved by actuators 66 operated by the control unit 500. In such way, an integrated laminar torsion-box type airfoil composite structure 2 comprising stiffening elements 8, 8', 8" and fastening elements 34 is provided.

Fig. 10f illustrates a device 510 according to one aspect of the invention. The control unit 500 of the production line PL described in Fig. 10a -10e can preferably comprise the device 510. The device 510 comprises a non-volatile memory NVM 520 which is a memory that can retain stored information even when the device 510 is not powered. The device 510 further comprises a processing unit 530 and a read/write memory 540. The NVM 520 comprises a first memory unit 550. A computer program (which can be of any type suitable for any operational data base) is stored in the first memory unit 550 for controlling the functionality of the device 510 as a part of e.g. the production line PL shown in Fig. 10a-10e.

Furthermore, the device 510 comprises a bus controller (not shown), a serial commu- nication port (not shown) providing a physical interface through which information transfers separately in two directions. The device 510 also comprises any suitable type of I/O module (not shown) providing input/output signal transfer, an A/D converter (not shown) for converting continuously varying signals from detectors (not shown) and other monitoring units (not shown) of the production line PL into binary code suit- able for the computer.

The device 510 also comprises an input/output unit (not shown) for adaption to time and date. The device 510 also comprises an event counter (not shown) for counting the number of event multiples that occur from independent events in the production line. Furthermore the device 510 includes interrupt units (not shown) associated with the computer for providing a multitasking performance and real time computing in said production line. The NVM 520 also includes a second memory unit 560 for external controlled operation. A data medium storing program P comprising driver routines adapted for drivers (not shown) and provided for operating the device 510 for performing the present method described herein. The data medium storing program P comprises routines for providing in a production line PL a torsion-box type airfoil composite structure 2. The data medium storing program P comprises a program code stored on a medium, which is readable on the device 510, for making the control unit 500 to perform the method steps of providing a forming tool, providing a skin component lay-up onto the forming tool, positioning stiffening element lay-ups and/or fastening element lay-ups on the skin component lay-up forming an integrated lay-up and co-curing the integrated lay- up.

The data medium storing program P further may be stored in a separate memory 570 and/or in a read/write memory 540. The data medium storing program P is in this embodiment stored in executable or compressed data format. It is to be understood that when the processing unit 530 is described to execute a specific function this involves that the processing unit 530 executes a certain part of the program stored in the separate memory 570 or a certain part of the program stored in the read/write memory 540.

The processing unit 530 is associated with a data port 555 for communication via a first data bus 515. The non-volatile memory NVM 520 is adapted for communication with the processing unit 530 via a second data bus 512. The separate memory 570 is adapted for communication with the processing unit 530 via a third data bus 511. The read/write memory 540 is adapted to communicate with the processing unit 530 via a fourth data bus 514. The data port 555 is preferably connectable to e.g. data links L501 , L502, L503, L504 and L505 of the production line PL shown in Fig. 10a-e.

When data is received by the data port 555, the data will be stored temporary in the second memory unit 560. After that the received data is temporary stored, the processing unit 530 will be ready to execute the program code, according to the above- mentioned procedure. Preferably, the signals (received by the data port 555) comprise information about operational status of the production line, such as operational status regarding the automated tape laying machine 64, the actuators 66 (hydrauls) and the autoclave 82. It could also be operational data regarding the positioning of the lay-ups 29, 62, 68 relatively the forming tools 40, 46, 60, 61 and the monitoring of such positioning. The received signals at the data port 555 can be used by the device 510 for controlling and monitoring a semi-automatic or automatic production line for manufacture of a laminar torsion-box type airfoil composite structure 2 in a cost-effective way. The signals received by the data port 555 can be used for automatically moving the composite structure 2 between the different operation positions in the production line by means of robot arms. The information is preferably measured by means of suitable sensor members arranged in each automatic apparatus of the production line. The information can also be manually fed to the control unit 500 via a suitable communica- tion device, such as a personal computer display.

Parts of the methods can also be executed by the device 510 by means of the processing unit 530, which processing unit 530 runs the data medium storing program P being stored in the separate memory 570 or the read/write memory 540. When the device 510 runs the data medium storing program P, the method steps disclosed herein will be executed.

A data medium storing program product (not shown) is also provided to perform the method steps when the data medium storing program (P) is run on the control unit 500.

The present invention is of course not in any way restricted to the preferred embodiments described above, but many possibilities to modifications, or combinations of the described embodiments, thereof should be apparent to a person with ordinary skill in the art without departing from the basic idea of the invention as being defined in the appended claims. The scarf joint does for example not have to be a stepped or a bevelled scarf joint but can be any type of scarf joint. A bevelled scarf joint is also called a plain scarf joint. The bevelled or stepped joining faces in a scarf joint may have an angle relative the plane of the skin component of 1-15 degrees, preferably 1 ,5-6 degrees. The torsion-box type airfoil composite structure according to the invention can form a part of any type of airfoil, such as for example an aircraft wing, fuselage, fin or nacelle. The torsion-box type airfoil composite structure according to the invention can comprise a plurality of stiffening elements and/or fastening elements. An adhesive can be used in the joints between the stiffening elements for further strength. Composite is defined as a cured resin comprising fiber reinforcement. The longitudinal direction regarding stiffening elements may be defined as a direction along the length of the stiffening elements. In the case where the torsion-box type airfoil composite structure is a part of an aircraft wing, the longitudinal direction (elongation) of the stiffening ele- ments is span wise along the wing. A direction transversal to the length of the stiffening elements, or transversal to the longitudinal direction, when the torsion-box type airfoil composite structure is a part of an aircraft wing, is a chord wise direction. The first forming tool can be a male tool and the second forming tool can be a female tool. The device used in the production line for performing the method steps disclosed herein can preferably be a computer.

Claims

1. A torsion-box type airfoil composite structure comprising a skin component (6, 206) and at least a first and a second stiffening element (8, 8', 208', 8", 208") arranged adjacent to each other longitudinally along the skin component (6, 206), wherein the at least first and second stiffening element (8, 8', 208', 8", 208") each comprise a flange member (14', 214', 14", 214") arranged in parallel with and bonded to the skin component (6, 206), wherein the second stiffening element (8, 8", 208") comprises a run-out area (10, 210), characterized in that the flange member (14', 214') of the first stiffening element (8, 8', 208') and the flange member (14", 214") of the second stiff- ening element (8, 8", 208") are joined at the run-out area (10, 210).

2. The structure according to claim 1 , wherein the skin component (6, 206), the first stiffening element (8, 8', 208') and the second stiffening element (8, 8", 208") are co- cured, forming an integrated structure.

3. The structure according to claim 1 or 2, wherein the flange member (14', 214') of the first stiffening element (8, 8', 208') and the flange member (14", 214") of the second stiffening element (8, 8", 208") are joined by a scarf joint at the run-out area (10, 210).

4. The structure according to claim 3, wherein the scarf joint is formed by correspondingly stepped faces (28', 28") of the flange members (14', 214', 14", 214") of the first stiffening element (8, 8', 208') and the second stiffening element (8, 8", 208") joined to each other.

5. The structure according to any of the preceding claims, wherein the flange member (14', 214') of the first stiffening element (8, 8', 208') comprises an extended part (24, 224) at the run-out area (10, 210) of the second stiffening element (8, 8", 208"), which part (24, 224) extends transversely to the longitudinal direction across the width of the flange member (14", 214") of the second stiffening element (8, 8", 208") and wherein the edge (26, 226) of the extended part (24, 224) facing the second stiffening element (8, 8", 208") is joined to the flange member (14", 214") of the second stiffening element (8, 8", 208") at least along a portion of the width of the flange member (14", 214") of the second stiffening element (8, 8", 208").

6. The structure according to any of the preceding claims, wherein the at least first and second stiffening element (8, 8', 8") are each configured with a T cross-section such that respective flange member (14', 14") of each stiffening element (8, 8', 8") comprises a first and a second flange (16', 16", 18', 18") and such that each stiffening element (8, 8', 8") comprises a web (12', 12"), extending at substantially right angle to the skin component (6).

7. The structure according to claim 6, wherein the stiffening elements (8, 8', 8") are each formed by two L-shaped profiles placed back-to-back with a blade (20', 20") in between.

8. The structure according to any of the claims 6-7, wherein the web (12") of the second stiffening element (8, 8") comprises a tapered portion (22) at the run-out area (10).

9. The structure according to any of the preceding claims, wherein the flange member (14', 214') of the first stiffening element (8, 8', 208') and the flange member (14", 214") of the second stiffening element (8, 8", 208") are longitudinally joined by a butt joint at least along a portion of the length of the second stiffening element (8, 8", 208").

10. The structure according to any of the preceding claims, wherein the stiffening elements (8, 8', 208', 8", 208") are stringers.

1 1. The structure according to any of the preceding claims, wherein the structure (2) comprises a fastening element (34) arranged within the run-out area (10, 210).

12. An aircraft wing comprising the structure (2) of any of the preceding claims.

13. A method for manufacturing a torsion-box type airfoil composite structure (2) comprising a skin component (6, 206) and at least a first and a second stiffening element (8, 8', 208', 8", 208") placed adjacent to each other longitudinally along the skin component (6, 206), wherein the at least first and second stiffening elements (8, 8', 208', 8", 208") each comprise a flange member (14', 214', 14", 214") arranged in parallel with and bonded to the skin component (6, 206), wherein the second stiffening element (8, 8", 208") comprises a run-out area (10, 210) and the flange member (14', 214') of the first stiffening element (8, 8', 208') and the flange member (14", 214") of the second stiffening element (8, 8", 208") are joined at the run-out area (10, 210), wherein the method comprises the steps of:

- providing a forming tool (40, 46, 60, 61);

- providing a skin component lay-up (62) onto the forming tool (60, 61);

- positioning stiffening element lay-ups (29) on the skin component lay-up (62) forming an integrated lay-up; and

- co-curing the integral lay-up.

14. The method according to claim 13, wherein the step of positioning the stiffening element lay-up (29) onto the skin component lay-up (62) is performed after forming said stiffening element lay-up (29) over a forming tool (40, 46).

15. The method according to claim 13 or 14, wherein the method further comprises the step of positioning fastening element lay-ups (68) on the skin component lay-up (62) prior to co-curing.

16. A production line for the production of a torsion-box type airfoil composite structure (2) comprising a skin component (6, 206) and at least a first and a second stiffening element (8, 8', 208', 8", 208") placed adjacent to each other longitudinally along the skin component (6, 206), wherein the at least first and second stiffening elements (8, 8', 208', 8", 208") each comprise a flange member (14', 214', 14", 214") arranged in parallel with and bonded to the skin component (6, 206), wherein the second stiffening element (8, 8", 208") comprises a run-out area (10, 210) and the flange member (14', 214') of the first stiffening element (8, 8', 208') and the flange member (14", 214") of the second stiffening element (8, 8", 208") are joined at the run-out area (10, 210), the production line comprising

-means for providing a forming tool (40, 46, 60, 61);

-means for providing a skin component lay-up (62) onto the forming tool (60, 61); -means for positioning stiffening element lay-ups (29) on the skin component lay-up (62) forming an integrated lay-up; and -means for co-curing the integrated lay-up.

17. The production line according to claim 16, wherein the production line further comprises means for positioning fastening element lay-ups (68) on the skin compo- nent lay-up (62).

18. A data medium storing program (P) for establishing in a production line according to claim 16 or 17 an automatic or semi-automatic manufacture of a torsion-box type airfoil composite structure (2), wherein said program (P) comprises a program code stored on a medium, which is readable on a device (510), to make a control unit (500) perform the method steps of:

-providing a forming tool (40, 46, 60, 61);

-providing a skin component lay-up onto the forming tool (60, 61);

-positioning a stiffening element lay-up (29) on the skin component lay-up (62) forming an integrated lay-up; and

-co-curing the integrated lay-up.

19. A data medium storing program product comprising a program code stored on a medium, which is readable on a device (510), to perform the method steps according to any of the claims 13 -15, when a data medium storing program (P) according to claim 18 is run on a control unit (500).