WO2014065718A1

WO2014065718A1 - An integrated curved structure and winglet strength enhancement

Info

Publication number: WO2014065718A1
Application number: PCT/SE2012/051126
Authority: WO
Inventors: Pontus Nordin; Anders RYDBOM; Tomas VÅNGELL
Original assignee: Saab Ab
Priority date: 2012-10-22
Filing date: 2012-10-22
Publication date: 2014-05-01
Also published as: EP2909009A4; EP2909009A1

Abstract

The present invention regards a structurally integrated curved structure (13, 13', 13'') comprising a curved transition region (15) being defined between a tip section (42) and a root section (11, 21), an upper and lower composite skin (27, 29) comprising an inner surface (45) fixed to a substructure (31) of said transition region (15), the substructure (31) and skins (27, 29) are, during use of the structure (13, 13', 13''), subjected to a bending moment (M), each skin (27, 29) of said transition region (15) comprises elongated reinforcement fiber-like elements. The lower composite skin (29) comprises reinforcing fibersand/or nano tubes(41, 43), which are oriented transverse to the inner surface (45). The invention also regard sa method for producing the structurally integrated curved structure (13, 13', 13'').

Description

An integrated curved structure and winglet strength enhancement

TECHNICAL FIELD

The present invention relates to structurally integrated curved structures according to the preamble of claim 1 and a method of manufacture of the structure according to claim 1 1 .

The invention relates to the aircraft industry and to aircraft maintenance engineering. The invention is not limited thereto, but can be used also in automotive engineering, trains, wind power blades etc.

BACKGROUND ART

Current technology is commonly based on assembled or co-cured carbon fiber reinforced plastic materials in laminate form. The current technology uses

mechanical fasteners and/or adhesive bonding of premanufactured fiber reinforced components, for strengthening a curved structure subjected to a bending moment.

There has been an issue for engineers to develop curved composite (cured resin comprising reinforcing elements) structures for strengthening the structures and make them efficient. Especially efforts to reduce fuel consumption of the aircraft are common or for make more efficient propellers and fans.

Bending moments are present in several curved structures of an aircraft during flight. One type of curved structure is a winglet having fiber reinforced composite skins. However, due to heavy loads on the curved transition region of the winglet, the region has to be reinforced by thicker composite laminate, by additional substructures etc. There are different types of winglets, such as retrofit winglets, wingtip fences, blended winglets, raked wingtips etc. The main geometrical parameters are the height, the taper ratio and the dihedral angle. If the dihedral angle is near constant, the winglet is almost planar and the transition region between the wing and the winglet root is curved. Blended winglets have a certain degree of sweep related to the rest of the wing. Raked wingtips comprise a higher degree of sweep and can be described as integrated wingtip extensions. There are S-shaped winglets and winglets having a constant radius curve rather than a sharp angle junction. There are winglets which also are curved inwardly over the wing. The winglets may be attached to the wing tip resulting in relatively sharp corners having a curvature of small radii. A common purpose of the major parts of winglets is to reduce the induced drag. In such way fuel can be saved. The winglet function is to relocate the wing tip vortex of the wing further outboard. However, during use, an inward pointing load generated by the winglet relocates the wing tip vortex, which produces a bending moment to the wing and winglet, especially high bending moments are due with respect to the curved transition region or to the sweep. Known winglets are disclosed in for example US2010/0019094, US 5407153, GB 2475523, US 2010/0155541 , US 5 348 253, US 1095952. GB 2475523 discloses a first skin, a second skin, an intermediate structure or substructure having contact portions connected to the inner surface of the skins. Interconnecting webs are integral arranged with the intermediate structure and extend between the inner surfaces of the skins. Such structure is provided to a winglet extending upwards from the end of the wing tip. GB 2475523 also discloses a method that separately forms the first skin, the second skin, the intermediate structure from composite materials. The skins and structure are co-cured in one cure step. The method comprises steps of stacking sheets of pre-impregnated fibers (so called pre-pregs) over mandrels and a second tool, the sheets have sections overlapping and forming the intermediate structure and the contact portion has a thicker portion. The substructure or ribs are formed by laying down a stack of six pre-impregnated carbon fiber sheets over aluminium alloy mandrels and closed by a lower skin tool in predetermined

configuration. The sheets can be laid down by an automatic tape laying machine (ATL). For proving strength, said lower skin being attached to the substructure by using rivets.

As the bending moment aspect also is critical for winglets in GB 2475523, radius fillers or deltoid fillers are placed in angles shaped by intermediate structure extending away from the inner surface of the skin. In such way the bending moment can be absorbed in a better way, especially within the area of the transition region of the winglet. There is an object to provide a fabrication of all types of structurally integrated curved structures, not only winglets, which during use absorb high bending loads. It could be structures such as flaps, landing gear door panels etc. It is desirable to make them all of low weight but at the same time of high strength, and to produce them in an efficient and cost-effective way of production.

There is an object to reduce fuel consumption of an aircraft comprising structural integrated curved structures.

There is an object to provide a smooth aerodynamic surface of a structurally integrated curved structure.

There is an object to decrease the structural weight of the aircraft. The use of winglets today involves strengthening the wing structure in said transition region because of the said increased wing bending moment.

There is an object to provide the transition region with a smooth aerodynamic surface.

There is an object to provide a wing still of low weight however designed to decrease induced drag by means of any type of winglet described above or other types, and solving the problem with bending and shear loads in the transition region.

There is an object to increase the efficiency of any type of structurally integrated curved structures producing a lifting force, such as wind power turbine blades, propellers, fans etc.

There is an object to provide a structurally integrated curved structure and method for producing the latter in a cost-effective way, and which is efficient to produce with short time in a production line. There is an object to reduce the number of details involved in the winglet attachments to the wing and also installation time and also to make automated manufacturing methods.

It is an object to strengthening an area of a winglet transition region subjected to interlaminar stresses due to bending of the winglet and transition region.

There is an object to provide a fabrication of structurally integrated curved structures, other than winglets, which during use absorb high bending loads. It could be structures as flaps, landing gear panels door etc.

The composite material used to form the substructure and skin may be of

unidirectional pre-impregnated composite material.

Alternatively, the material used can be pre-impregnated woven fabric.

Alternatively, the material used can be dry woven fabric, in a second step injected or impregnated with resin.

Alternatively, the material used can be dry woven fabric interleaved with resin film.

SUMMARY OF THE INVENTION This has been achieved by the structure defined in the introduction and being characterized by the features of the characterizing part of claim 1 .

In such way is a structural integrated curved structure achieved which has high strength and low weight. Saving weight means that also fuel is saved. No rivets or stiffening plates have to be used for strengthening e.g. a winglet transition region. Thereby is achieved a smooth aerodynamic surface of the transition region, which also saves fuel during flight. Preferably, also the upper composite skin comprises reinforcing fibers or/and nano-sized inclusions, which are oriented transverse to the inner surface. Suitably, the transverse oriented reinforcement structure comprises CNTs (carbon nano tubes), which are grown in radial direction directly on individual large carbon fibers. The transverse oriented fibers may be three dimensional carbon fiber fabrics (tailor made three dimensional fabrics and arrangement of stitched fabric assemblies to be adapted to the actual design). Laminate materials may be based on pre-preg tape such as unidirectional pre-impregnated fiber plies, fibers being of woven carbon fiber pre-preg fabrics or glass, Kevlar, spectra pre-preg tapes and fabrics etc.

Preferably, the transition region comprises transverse oriented elongated reinforcing fiber-like elements, such as reinforcing fibers and/or carbon nano tubes.

Thereby bending and shear loads can be taken up by means of a curved transition region between the winglet or winglet tip section and the wing or wing section. In such way a transition region of low weight and high strength is achieved.

Suitably, the skins of a wing root of the wing section also comprise transverse oriented elongated reinforcement fiber-like elements (such as reinforcing fibers and/or carbon nano tubes) for taking up high bending loads. In such way is achieved that that also other sections of the wing being subject for bending moment due to the application of a winglet at the wing tip can be

strengthening, and the structure having low weight.

Preferably, the upper composite skin comprises transverse reinforcing fibers.

Preferably, the reinforced composite skin comprises at least two plies.

Thereby a laminate stack can be built by means of an ATL (automatic tape laying apparatus automatically laying pre-preg stacks onto each other)-apparatus which is cost-effective. Each skin ply may comprise large fibers, such as carbon fibers, graphite fibers and/or carbon nano tubes oriented in different directions. For example, a ply with span wise oriented carbon fibers could be laid onto and adjacent a ply with chord wise oriented carbon fibers and upon this one a further ply having 45 degrees oriented fibers relative the span wise direction. Each ply can also comprise large fibers and/or carbon nano tubes (generic term for such large fibers and CNTs: elongated reinforcement fiber-like elements) oriented transverse relative the surface of the laminate stack. In such way the strength transverse the skin surface is achieved within the area of the transition region and a low weight structure can be achieved. The fibers used for the transverse direction can be of the type used for span wise oriented fibers, which is cost-effective.

Suitably, the transverse oriented elongated reinforcement fiber-like elements are carbon fibers.

In such way a cost-effective production of the wing is achieved.

Preferably, the transverse oriented elongated reinforcement fiber-like elements are CNTs.

Thereby is achieved a strengthening of the laminate by a simple way of production within the curved transition region of the structure or the sweeping region of a blended winglet. Preferably, also nano fibers, nano wires or other nano filament structures can be used. The application of transverse oriented nano filaments relative the skin surface can be achieved in different ways. For example, growth of nano tubes onto large fibers, which large fibers extends in the plane of the skin. In such case, some of the nano filaments are oriented transverse. Other examples may include a skin having all nano filaments oriented transverse the skin surface. Suitably, the composite skin comprises at least two plies and aligned transverse oriented CNTs being positioned between the plies for strengthening a bond between the plies in a transverse direction within the transition region.

Suitably, the third fiber material of the three-dimensional-engineered fiber materials can include be aligned transverse oriented carbon nano tubes or nano wires in- between the plies.

In such way the strength within the curved transition region of the structure being achieved in a cost-effective way. An entire winglet may be produced with strengthening nano filament structures between and/or inside the plies, which structures or part of structures are oriented transverse the plane of the skin.

Preferably, the substructure comprises transverse oriented reinforcing fibers being transverse oriented relative the inner surface.

In such way is achieved that also the ribs and spar caps and beams and spars can be manufactures with high strength in a cost-effective way, especially within the area of the transition region.

Suitably, the inner surface is fixed to the substructure via an adhesive comprising a nano filament structure, preferably within the curved transition region.

In such way is achieved an enhanced bending load absorbing functionality of the transition region of the curved structure. Alternatively, the substructure is a honeycomb material and the skins are adhered to the honeycomb outer surface. Alternatively, nano-engineered adhesives for bonding of pre-cured spar segments (spars, spar caps etc.) to pre-cured or "wet" laid up skins or plies. This is also solved by a winglet profile formed to be fixed between a winglet tip section and a wing section, the winglet transition region being curved and having an upper and lower composite skin, wherein at least the lower reinforced composite skin comprises reinforcing fibers, which are oriented transverse to the inner surface. In such way is achieved that a winglet transition profile for coupling a winglet tip to a wing tip is achieved. In such way, due to the strength of the winglet transition profile achieved without any reinforcing plates and rivets disturbing the airflow, the aircraft will use less fuel. The winglet transition profile is preferably curved after a line being transverse to a direction corresponding with the flight direction (during use of the article), which line follows the elongation of the winglet.

This has been achieved by the method defined in the introduction and characterized by the steps of claim 1 1 . In such way is achieved a cost-effective way to produce a structural integrated curved structure and/or winglet transition profile. The method provides an article that involves high strength without incorporating any heavy or laminar airflow disturbing details within the area of the curved transition region. The method further is cost- effective since it can be performed efficiently in an already set production line. The method may use known production tools and just minor adjustments of production equipment have to be made. The use of three-dimensional-engineered CFRP

(Carbon-fiber-reinforced polymer) materials will allow the design and fabrication of structurally integrated curved structures such as winglets and similar curved structures. The application of such curved structures, separately or in combination, will allow design solutions such as winglets, antenna fairings, curved landing gear doors, etc.

The method provides a curved structure (article) that involves high strength (having no heavy or turbulence generating disturbing details) within the curved transition region for counteracting the bending moment of the structure. The design regards structures which are subjected to a bending moment, and which structures are curved in a transition region. For example, a curved antenna fairing having an interial structure bonded to a curved cover housing is also affected by bending loads and also comprises a corresponding transition region. The bonding between the skin inner surface and the ribs, are preferably made with an adhesive comprising transverse oriented fibers.

The use of composite materials having the improved intralaminar and interlaminar strength and stiffness will make it possible to design a structural profile, such as wing box or wing profile or winglet profile, in an integrated way allowing co-curing of skin panels with the substructure (ribs, spares, spar caps, beams etc.) without the use of mechanical fasteners. Suitably, the composite skin comprises at least two plies with aligned transverse oriented carbon nano tubes being positioned between the plies for strengthening the joint between the plies in a transverse direction relative the ply plane and within the joint region. Preferably, the elongated reinforcement fiber-like elements exhibit through-thickness reinforcement with extension partly or entirely through one, more than one or all plies of the structural composite parts within the joint region. Suitably, the three-dimensional-engineered fiber material can include aligned transverse oriented carbon nano tubes or nano wires in-between the plies for further improvement of the peel strength.

Preferably, the nano-sized elongated reinforcement fiber-like elements (CNT, nano fibre, nano multi wall filament, nano double wall filament, nano wire etc.) has a length of 0, 1 mm up to 3,0 mm for achieving improved strength between rib foot and skin. This is suitable for a common pre-preg ply having a thickness of 0, 125 mm used in the production of aircrafts. If leaning (relative laminate plane) nano-sized elongated reinforcement fiber-like elements are used, the length preferably can be longer. The definition of nano is an element having at least one dimension not more than 200 nm. 1 nm (nanometre) is defined as 10^"9 metre (0,000 000 001 meter). Preferably, the diameter of a multiwall nano tube is 15-35 nm, suitably 18-22 nm. Suitably, the diameter of a single wall nano tube is 1 ,2-1 ,7 nm, preferably 1 ,35-1 ,45 nm. In such way the strength within the joint region of the structure being achieved in a cost-effective way.

Alternatively, the elongated reinforcement fiber-like elements exhibit through- thickness reinforcement with extension through the at least on ply with a

measurement of about from 10 pm to 10 mm, preferably 127 pm to 7 mm. Preferably, in case of using CNTs, the through-thickness reinforcement may have an extension up to 1 ,5 mm into the skin laminate and into the flange of the rib foot respectively.

Suitably, the curing tool is a hot drape tool.

In that way is achieved that a cost-effective method is achieved. This is achieved by that a wing let manufacturer achieves dimensional control inside the hollow structure at the same time as faster curing cycles are used. The use of laid-up pre-preg forming the skins gives the possibility to prepare the pre-preg tape, not only with in- plane arranged elongated reinforcement fiber-like elements but also with transverse arranged elongated reinforcement fiber-like elements. The skins are preferably used for the curved transition region of the structure for absorbing the bending loads of the structure during flight.

Preferably, the reinforced composite skin of the curved transition region is single curved or double curved. Preferably, the cure tool is an injection tool (RTM-resin transfer moulding).

Suitably, the cure tool is an autoclave apparatus.

Alternatively, the cure tool is an out-of autoclave apparatus.

The laminate materials may be achieved by resin transfer moulding (RTM) in an automated operation that combines compression, moulding, and transfer moulding processes. In such way has been achieved a good surface finish and a dimensional stability in a cost-efficient way. The mould is loaded with layers of dry fibers and the resin being injected or drawn into the mould. Some of the elongated reinforcement fiber-like elements are oriented transverse to the laminate surface. The resin is heated and pressure is applied. After curing the curved article, such as the present winglet or winglet transition profile, is removed, trimmed, ID-marked and stored. Suitably, the method comprises the further step of applying aligned transverse oriented CNTs in position between at least two plies for strengthening the bond between the plies in transverse direction, preferably within the curved transition region. In such way improved strength is achieved within the transition region of the structure. An entire winglet may be produced with transverse oriented strengthening nano filament structures between and/or inside the plies, which nano structures or part of nano structures are oriented transverse the plane of the reinforced composite skin. Preferably, the method comprises the further step of fixing the inner surface of the single curved/double curved composite skin to the substructure via an adhesive comprising a nano filament structure, preferably within the transition region.

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 a straight wing;

Fig. 2 illustrates a wing comprising a winglet;

Fig. 3 illustrates, in a perspective view, a wing comprising a winglet;

Figs. 4a-4d illustrate examples of winglets;

Figs. 5a-5c are illustrations of a retrofit (or detachable) winglet;

Figs. 6a-6c are illustrations of a blended winglet of the type shown in Fig. 4c;

Figs. 7a-7d are illustrations of a blended winglet of the type shown in Fig. 4c according to an embodiment of the present invention;

Figs. 8a-8f illustrate a winglet comprising a transition region and a distal wingtip end; Fig. 8g illustrates a substructure bonding to a skin inner surface;

Fig. 8h illustrates a three-dimensional-engineered carbon-fiber-reinforced materials;

Figs. 9a-9e illustrate a method for manufacture of a curved airfoil structure;

Figs. 10a and 10c illustrate a transition region of another curved article of an aircraft;

Fig. 10 b illustrates prior art transition region;

Figs. 1 1 a and 1 1 b illustrate different types of moulding tools; and

Fig. 12 illustrates an autoclave apparatus.

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 may be deleted from the drawings. Also, the illustrative drawings show fiber structures of different types, being illustrated extremely exaggerated and schematically for the understanding of the invention. The nano structures are illustrated exaggerated in the figures also for the sake of understanding of the orientation and the alignment of the nano filaments. The joint region is defined as the region where a composite skin inner surface and a sub-structure are joined to each other within the transition region.

Fig. 1 schematically illustrates an ordinary straight wing 1 of an aircraft 2 seen from front in a direction corresponding with the direction of airflow or direction of flight. At the wing's 1 wingtip 3 a vortex 5 is formed due to the difference in air pressure between upper 7 and lower 9 surfaces (under pressure U and over pressure 0) of the wing 1 . Air is caused to escape around the wingtip 3 which will reduce the lift and increase the drag of the aircraft 2.

Fig. 2 schematically illustrates a wing 1 1 comprising a winglet 13. With the winglet 13 mounted to the wing 1 1 of the aircraft, the outboard sections of the wing 1 1 will produce more lift. During use, an inward pointing load is generated by that the winglet 13 relocates the wingtip vortex 5 and produces a bending moment M to the wing 1 1 and winglet 13, especially in an area called the curved transition region 15 defined between the wing 1 1 and the winglet 13. Also the bending moment M' adjacent the wing's 1 1 root 17 increases significantly when the winglets 13 are used. The winglet 13 thus produces an inward load as it diverts the tip vortex 5 outward. At the same time, the wingtip will generate more lift which also increases the bending moments M, M'.

Fig. 3 schematically illustrates in a perspective view a wing 1 1 comprising a winglet 13. The Fig. 3 illustrates three axes X, Y, Z. The span wise direction essentially extends along the Y-axis. The chordwise direction of the wing 1 1 essentially corresponds with the X-axis. The third axis is the Z-axis, which is orthogonal to the X- and Y-axes. The winglet 13 and transition region 15 exhibits a curvature, which curves in a direction towards the Z-axis.

In Figs. 4a-4d different types of winglets 13 are schematically illustrated in

perspective. The main geometrical parameters are the height, the taper ratio and the dihedral angle, but the aerodynamic principle is almost the same for all winglets 13. The common purpose of the winglets 13 is to reduce the induced drag. In such way fuel can be saved. Fig. 4a illustrates a blended winglet 13 with a high degree of sweep relative the rest of the wing 1 1 . The dihedral angle is near constant and the winglet 13 is almost planar in a direction extending with a sharp angle relative to the Z-direction and a transition region 15 between the wing 1 1 and the winglet 13 root 17 being curved. Fig. 4b illustrates a retrofit winglet 13 being mounted onto a wingtip 21 . The retrofit winglet 13 is thus attached to the wingtip 21 in a way resulting in relatively sharp corners 23 of the transition region 15 having a curvature of small radii. Fig. 4c illustrates a blended winglet 13 with a low degree of sweep relative the rest of the wing 1 1 . The blended winglet 13 exhibits a smooth curve which reduces shock interference between wing 1 1 and winglet 13. The transition region 15 is formed with a nearly constant radius of curvature of sweep. Fig. 4d illustrates a wingtip fence 13', which also functions to relocate the wingtip vortex of the wing 1 1 further outboard. This kind of winglet 13' also comprises a transition region 15. However, this region exhibits a small radius of curvature of structure. During use, an inward pointing load is generated by the wingtip fence 13' (for relocating the wing tip vortex) which produces a bending moment to the curved transition region 15.

Fig. 5a schematically illustrates a side view of a cut-away of the retrofit or detachable winglet 13 shown in Fig. 4b. A winglet root 19 of the winglet 13 is attached to the wing's 1 1 wingtip 21 . Bolts 25 mechanically connect the winglet 13 to the wingtip 21 . An upper 27 and lower 29 skin are attached to a substructure 31 of the winglet 13 and wing 1 1 and also of the transition region 15. For reinforcing the transition region 15 in view of said bending moment M, which will occur during use of the winglet 13, a structural rib 33 is affixed to the substructure 31 and designed with a curvature.

However, in order to minimize added weight (caused by further reinforcing structures) to the transition region 15, the upper 27 and lower 29 skins of graphite epoxy are provided with carbon nano tubes (not shown). The carbon nano tubes are oriented in a direction orthogonally to the upper 27 and lower 29 skin surface. The orthogonally oriented carbon nano tubes of upper skin 27 take up forces acting on the upper skin 27 (inboard skin 27'), which forces will compress the skin 27 in a direction parallel with the plane of the upper skin 27. The orthogonally oriented carbon nano tubes of lower skin 29 take up forces acting on the lower skin 29 (outboard skin 29') with tension in a direction parallel with the plane of the lower skin 29. The winglet 13 is attached to wing tip ribs 34 by means of the bolts 25. A spar cap 35 and a root rib 37 of graphite epoxy are provided for acting with bending stiffness and strength to the transition region 15 at a minimum weight. The inboard 27' and outboard 29' skins are made of laid up CFRP or another structural composite material. The bending stiffness being provided by the orthogonally oriented carbon nano tubes and by the fiber reinforced plastic substructure 31 . Shear forces within the skins 27, 29 will be counteracted by the application of the transverse (relative the plane or extension of skins) oriented carbon nano tubes.

Stress and stiffness loads are thus absorbed by the substructure 31 , the structural ribs 33, the bolts 25 and the carbon nano tubes CNT's oriented transverse to the skin 27, 29 surface. Thereby the winglet root 19 will have a load distribution also taken care of by the skins 27, 29. Fig. 5b schematically illustrates from above the section A- A in Fig. 5a of the transition region 15. The structural wing tip ribs 34 are formed by middle and aft wing rib 39 and the ribs of the winglet 13 are attached via the bolts (not shown) to said wing ribs 39. In Fig. 5c is a further view illustrating the winglet 13 in Fig. 5a from the side.

Fig. 6a schematically illustrates a blended winglet 12 of the type shown in Fig. 4c and is made according to prior art. The blended winglet 12 exhibits a low degree of sweep and a smooth curve which reduces shock interference in the transition region 15. The upper and lower skins of the transition region 15 is made of epoxy or other resin pre-preg plies being reinforced by carbon or other fibers (e.g. cyanate ester) extending in a direction parallel with the skin extension. Such blended winglet 12 according to prior art must be re-inforced within the curved transition region 15 because of the increased bending moment of the wingtip 3. Such prior art

reinforcement is often made by thicker skin within the transition region 15 area and/or more rigid wing ribs within said transition region 15. Fig. 6b schematically illustrates in a closer view from above a section of the prior art skin regarding the fiber orientation. As can be seen, each ply P comprises fibers f having a specific orientation in a common plane so that when all plies P are laid up onto each other, the laminate skin exhibits fiber directions of 0 degrees, 90 degrees, +/-45 degrees parallel with the skin plane. Fig. 6c schematically illustrates a cross section of a skin laminate according to prior art within the transition region 15. As can be seen each ply P comprises fibers f having an orientation along the plane of the skin within the transition region 15.

Fig. 7a schematically illustrates a blended winglet 13 of the type shown in Fig. 4c according to an embodiment of the present invention. In Fig. 7a is disclosed that the wing skins (upper 27 and lower 29, only upper 27 being shown) within the transition region 15 between the wing 1 1 and the winglet 13 also comprises reinforcement (carbon) fibers 41 being oriented transverse to the skin plane extension EX. A section of the skin is shown in Fig. 7b from above (within the curved transition region 15) with a more detailed fiber orientation. As is shown, the laminate skin 27 exhibits fiber directions in 0 degrees, 90 degrees, +/-45 degrees as the skin shown in Fig. 6, but also carbon fibers 41 exhibiting a transverse orientation relative the skin 27, 29 plane. Fig. 7c schematically illustrates a cross section of the skin 27 laminate in Fig. 7b. As can be seen in Fig. 7c, each ply P comprises fibers f having an orientation parallel with the plane extension EX of the skin 27 within the transition region 15 and also comprises transverse fibers 41 having an orientation transverse to the skin 27 plane extension EX. In such way is a winglet 13 achieved which has high strength and low weight. Saving weight means that also fuel is saved. No further material has to be used for strengthening the winglet 13 for mounting the latter to the wing 1 1 and therefore no rivet heads will make any turbulence, instead is achieved a smooth aerodynamic surface of the transition region 15, which also saves fuel during flight. Fig. 7d schematically illustrates how the bending moment M produces a bending load upon a curved structure and the forces acting on the transition region 15. The upper skin 27 (inboard skin) is affected by forces pressing (compressing in a direction corresponding with the plane extension EX) the skin 27 whereby the skin 27 has a tendency to delaminate and the forces will make a tension F1 of the laminate in a direction orthogonally to the skin 27 surface and in a direction from the latter. The lower skin 29 (outboard skin) is affected by forces making a traction action upon the skin 29 whereby the skin 29 (and the plies of the skin 29) has a tendency to become thinner by compressing forces F2. By adding the carbon fibers 41 (see Fig. 7c) exhibiting a transverse orientation relative the respective skins 27, 29, the said forces (tension F1 /compressing F2) are possible to counteract and the laminate of the skins 27, 29 will maintain its strength. The substructure 31 is positioned between the skins 27, 29. No additional nor superfluous strengthening structures such as ribs, spar caps, stringers etc. have to be mounted within the transition region 15 for taking up the bending moment M. The substructure 31 can be made with less reinforcing material due to the new property of the skins 27, 29. Each ply thus comprises portions of large fibers 41 oriented transverse relative the surface of the skin 27, 29 and the laminate stack plane extension EX. In such way completary strength in a direction transverse the skin 27, 29 surface (or plane extension EX) being achieved within the transition region 15 and thereby a low weight structure can be achieved.

Figs. 8a-8f schematically illustrate a winglet 13 comprising a transition region 15 and a distal wingtip 21 end. The transition region 15 is defined between a winglet tip section 42 of the winglet 13 and a wingtip 21 of the wing 1 1 . The wing 1 1 extends along an Y-axis. The wing's 1 1 wing chord extends along an X-axis. These axes are perpendicular to a Z-axis. The transition region 15 is curved, wherein the winglet tip section 42 declines towards the Z-axis. The winglet's 13 transition region 15 comprises an upper 27 and lower 29 composite skin. The respective skin 27, 29 exhibits an inner surface 45 being fixed to ribs 47 forming a substructure 31 of the winglet 13. Each skin 27, 29 comprises a reinforcing fiber structure in the form of carbon fibers. The wingtip 21 end is adapted for attachment of the winglet 13, wherein the substructure 31 of the winglet 13 being pushed into the wing 1 1 structure. A cross section A-A is shown in Fig. 8b (the ribs 47 have been co-cured with upper 27 and lower 29 skins in one curing cycle process). Upper 49 and lower 50 flanges of the ribs 47 are bonded with the inner surfaces 45 of the respective upper 27 and lower 29 skins. In Fig. 8b is shown an area defined by a dashed and dotted circle line, which area is illustrated in Figs. 8c-8f by examples of the

reinforcing fiber structure of the lower 29 composite skin comprising reinforcing fibers 43', 43", 43'", 43"" oriented transverse to the inner surface 45 of the skin 29. Fig. 8c illustrates lower flange 50 adhered to the inner surface 45 of the winglet 13 skin 29 of a transition region 15. An adhesive 51 bonds the flange 50 with the inner surface 45. The adhesive 51 comprises an epoxy and carbon nano tubes CNT's 43' having an orientation perpendicular to the plane of the inner surface 45. The reinforcing fiber structure (not shown) of the lower composite skin 29 thus comprises the adhesive 51 including reinforcing fibers 43', which are oriented transverse to the inner surface 45. The adhesive 51 is defined as a portion of the skin 29. Fig. 8d illustrates a further embodiment. Plies P of pre-pregs (beforehand resin impregnated fiber sheets) have been stacked onto each other. The plies P are laid by means of an ATL-apparatus (not shown). The transverse oriented reinforcing fibers are in this embodiment elongated nano wires 43", which are arranged in such way that a major part of them joins two adjacent plies P, wherein the interface bond between the plies P is strengthened in transverse direction by means of the transverse nano wires 43". A further embodiment is shown in Fig. 8e. In this example, transverse nano fibers 43"' are disposed within each ply P in a transverse direction relative the ply P extension. A further embodiment is disclosed in Fig. 8f wherein some of carbon nano tubes 43"" (of a growth of aligned carbon nano tubes NF onto large fibers f of the plies P) are oriented transverse to the skin 29 extension. The large fibers f extend along the plane of the skin 29. Thereby is achieved a strengthening of the laminate skin 29 and/or the bonding between the substructure 31 and the skin 29 by a cost-effective manner of production strengthening the transition region 15 of the winglet 13 or the sweeping region of a blended winglet. Fig. 8g schematically illustrates a flange 50 of a rib bonded to an inner surface 45 of a skin 29 within a curved section of an airframe subjected to bending moments. Also the flange 50 of the substructure comprises transverse oriented nano tubes. Fig. 8h illustrates a further embodiment. A three-dimensional-engineered carbon- fiber-reinforced material 3D is applied to the plies P for bonding them together.

Thereby is achieved a strengthening of the laminate skin 29.

Figs. 9a-9e schematically illustrate a method for manufacture of a winglet. A semi- cured skin 55 of laid up pre-pegs made by an ATL apparatus 57 as shown in Fig. 9a. A moulding tool 58 provided for forming the winglet to a proper form being disclosed in Fig. 9b. The semi-cured skin 55 is applied in the tool's 58 bottom forming part and well defined substructure semi-cured ribs 59 are placed onto and fixed to the skin 55 inner surface in contact with interior forming fly away tool parts (not shown). An upper semi-cured skin 55' is applied onto the substructure 31 .

The design in this case regards dual-skin structures (upper and lower skin) having a sub-structure 31 (supporting structure in the form of spars, ribs and/or beams) there between. Ribs 59 are chordwise structures (extending in the direction of flight) connecting together the upper 55' and lower 55 skin (also called inboard and outboard skin within the area of the finished winglet 13). The function of the ribs is to maintain the aerodynamic profile of the wing and winglet. They also transfer air pressure (over-/under pressure) from the skins 27, 29 to the spar caps 35 and they diffuse locally concentrated load inputs and they function to redistribute wing bend loads and winglet root bend moment loads

In Fig. 9c is shown an upper part of the moulding tool 58 for forming the winglet 13. The winglet 13 thus comprises the curved transition region 15 having an upper 27 and lower 29 composite skin, each comprising an inner surface fixed to the substructure 31 , wherein each composite skin 27, 29 comprises a reinforcing fiber structure at least partly having a fiber orientation being transverse to the inner surface 45. Curing of the winglet 13 is made by a co-curing process (not shown) and the winglet 13 is removed from the moulding tool 58 after curing, as is shown in Fig. 9d. In Fig. 9e is shown the winglet 13' mounted to a wing 1 1 . The winglet 13' comprises a curved transition region 15 (comprising the upper 27 and lower 29 composite skin, each exhibiting a reinforcing fiber structure which at least partly having a fiber orientation being transverse to the inner surface 45 shown in Fig. 9c). Fig. 10a illustrates a transition region 15 of another structurally integrated curved structure of air stream dividing flap 13" of an aircraft, which curved structure takes up a bending moment during the use of the structure. Tension forces strive for making the lower skin 29 thinner and break up the bond between the substructure 31 and the skin 29 inner surface. Compression forces C strive for making the skin thicker and delaminating the skin 29. The curved flap 13" article being affected by a bending moment M, during its use and motion through the air stream.

According to prior art and Fig 10b, a curved flap must be built with heavy

strengthening structures, which also sometimes includes fasteners that protrude from the aerodynamic surface. A curved flap having a weak substructure according to prior art will over time delaminate as shown in Fig. 10b.

Fig 10c schematically illustrates transverse fibers 43 according to a further example of the actual flap 13" in Fig. 10a according to a further embodiment of the present invention. Some structural fibers of the transition region 15, such as transversal fibers 43, are arranged within the lower skin 29 for counteracting the tension of the skin 29 caused by moment M. Fig. 1 1 a schematically illustrates a moulding tool which is a hot drape tool 58'. Fig. 1 1 b schematically illustrates a moulding tool which is an infusion tool 58". The laminate materials are in this example produced by resin transfer moulding (RTM) in an automated operation that combines compression, moulding, and transfer moulding processes. The moulding tool 58" is loaded with layers (not shown) of fibers and the resin being injected into the tool. Some of the fibers are oriented transverse to the laminate surface. The resin is heated and pressure is applied. After curing, the structural integrated curved structure (not shown) is removed, trimmed, painted, ID-marked and stored. Fig. 12 illustrates an autoclave apparatus AC. The lay-up LU comprising the elongated reinforcement fiber-like elements (oriented transverse relative the surface of the laminate stack) is placed onto male tools MT. The lay-up is enclosed by a vacuum bag VA and being evacuated. Subsequently the lay-up and vacuum bag moved into the autoclave apparatus for pressurizing the lay-up under application of heat for curing the lay-up forming the winglet.

The method provides a curved article that involves high strength (having no heavy or turbulence generating disturbing details) within the transition region for counteracting the bending moment of the article. The use of three-dimensional-engineered e.g. CFRP (Carbon-fiber-reinforced polymer) materials will allow the design and fabrication of structurally integrated curved structures. The application of such curved structures, separately or in combination, will allow design solutions, such as winglets, antenna fairings, curved landing gear doors, etc. The design regards structures which are subjected to a bending moment, and which structures are curved in a transition region. For example, a curved antenna fairing having a structural interior bonded with curved cover housing is also affected by bending loads and also comprises a corresponding transition region. The bonding between the skin inner surface and the ribs are made of adhesive comprising transverse oriented fibers, carbon nano tubes or similar reinforcement. 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 defined in the appended claims. Examples of suitably composite materials include carbon fibers, aramid fibers, glass fibers, combinations of carbon and glass fibers or combinations of carbon, aramid and glass fibers. The skins and the substructure may be provided by laminating multiple sheets or layers having the fibers of respective sheet oriented in different directions, however having at least within the area of the transition region of the winglet, fibers, nano tubes or similar reinforcements oriented transverse the sheets major extension. The invention is not limited to aircraft. Also other aerial vehicles can be actual for implementation of the invention. Instead of winglet, the word integrated curved structure may be used. Such structure may be used also for applications not related to aerial vehicles, but wind mills, high speed trains etc. The design could also regard a curved structure having outboard and inboard skins arranged against a distance holding structural honeycomb.

Claims

1 . A structurally integrated curved structure (13, 13', 13") comprising:

-a curved transition region (15) being defined between a tip section (42) and a root section (1 1 , 21 );

-an upper and lower composite skin (27, 29) comprising an inner surface (45) fixed to a substructure (31 ) of said transition region (15);

-the substructure (31 ) and skins (27, 29) are, during use of the structure (13, 13', 13"), subjected to a bending moment (M);

-each skin (27, 29) of said transition region (15) comprises a reinforcing fiber structure;

characterized by that

-the lower composite skin (29) comprises elongated reinforcement fiber-like elements (41 , 43), which are oriented transverse to the inner surface (45).

2. The structure according to claim 1 , wherein the transition region (15) comprises transverse oriented elongated reinforcement fiber-like elements (43', 43", 43"', 43"").

3. The structure according to claim 1 or 2, wherein the upper composite skin (27) comprises transverse elongated reinforcement fiber-like elements (41 , 43).

4. The structure according to any of claim 1 to 3, wherein the composite skin (27, 29) comprises at least two plies (P).

5. The structure according to any of the preceding claims, wherein the transverse oriented elongated reinforcement fiber-like elements are carbon fibers (41 ).

6. The structure according to any of the preceding claims, wherein the transverse oriented elongated reinforcement fiber-like elements are CNT (43).

7. The structure according to any of the preceding claims, wherein the composite skin (27, 29) comprises at least two plies (P) and aligned transverse oriented CNTs (43") being positioned between the plies (P) for strengthening a bonding between the plies (P) in a transverse direction within the transition region (15).

8. The structure according to any of the preceding claims, wherein the substructure (31 ) comprises transverse oriented elongated reinforcement fiber-like elements (41 , 43) being transverse oriented relative the inner surface (45).

9. The structure according to any of the preceding claims, wherein the inner surface (45) is fixed to the substructure (31 ) via an adhesive (51 ) comprising a nano filament structure (43'), preferably within the transition region (15).

10. A winglet profile formed to be fixed between a winglet tip section (42) and a wing section (1 1 ), the winglet transition region (15) being curved and having an upper and lower composite skin (27', 29'), wherein at least the lower composite skin (29') comprises elongated reinforcement fiber-like elements (41 , 43), which are oriented transverse to the inner surface (45) of the skin (29').

1 1 . A method for manufacture of a structurally integrated curved structure (13, 13', 13"), comprising a curved transition region (15) having an upper (27) and lower (29) composite skin, each comprising an inner surface (45) fixed to a substructure (31 ), wherein each composite skin (27, 29) comprises elongated reinforcement fiber-like elements (41 , 43) at least partly having orientation being transverse to the inner surface (45); the method comprises the steps:

-providing a moulding tool (58', 58") for forming the structure (13, 13', 13");

-applying uncured or semi-cured resin and the reinforcing fiber (f, 41 , 43) structure to the tool (58', 58") for providing the upper and lower composite skin (27, 29) and the substructure (31 ), the substructure (31 ) being in contact with the inner surface (45); -curing the structurally integrated curved structure (13, 13', 13");

-removing the structure (13, 13', 13") from the moulding tool (58', 58").

12. The method according to claim 1 1 , wherein the moulding tool is a hot drape tool (58').

13. The method according to claim 1 1 , wherein the moulding tool is an infusion tool (58").

14. The method according to any of claims 1 1 or 13, the method comprises the further step of applying aligned transverse oriented CNT (43") in position between at least two plies (P) for strengthening the bond between the plies (P) in transverse direction, preferably within the transition region (15).

15. The method according to any of claims 1 1 to 14, the method comprises the further step of fixing the inner surface (45) to the substructure (31 ) via an adhesive (51 ) comprising a nano filament structure (43'), preferably within the transition region.