FUSELAGE STRUCTURE OF AN AIRCRAFT AND METHOD FOR MANUFACTURING THE SAME

Abstract

A fuselage structure of an aircraft includes a fuselage skin, and a plurality of frame elements spaced apart from one another in a direction parallel to the aircraft longitudinal axis for supporting the fuselage skin. The fuselage skin includes a plurality of interconnected fiber-reinforced composite skin panels that extend between each pair of frame elements and are connected thereto. The composite skin panels further comprise a stiffener integrally formed in each composite skin panel. A method for manufacturing the fuselage skin. The composite skin panels may be interconnected and/or connected to a frame element through an induction welded connection.

Claims

1-26. (canceled)

27. A fuselage structure of an aircraft comprising: a fuselage skin, extending along a longitudinal axis of the aircraft and enclosing an inner space, further having an inner surface facing the inner space, and a plurality of frame elements spaced apart from one another in a direction parallel to the aircraft longitudinal axis and extending in a circumferential direction along the inner surface of the fuselage skin to support the fuselage skin, wherein the fuselage skin (2) comprises a plurality of fiber-reinforced composite skin panels that are interconnected via second wall parts of said composite skin panels, and that extend between each pair of frame elements, wherein first wall parts of a composite skin panel are connected with first wall parts of a frame element, wherein the composite skin panels further comprise a stiffener integrally formed in each composite skin panel and extending radially inwards from the inner surface, wherein the stiffeners extend in a direction parallel to the aircraft longitudinal axis, wherein the first wall parts of the composite skin panels are located more radially inwards than the first wall parts of the frame elements to which they are joined, and wherein at least some of the first wall parts of the composite skin panels and the frame elements and/or at least some of the second wall parts of the composite skin panels are joined through an induction welded connection.

28. The fuselage structure according to claim 27, wherein the frame elements have an I-shaped or H-shaped cross-section, and the first wall part of the frame element comprises a flange of the I- or H-shaped frame element.

29. The fuselage structure according to claim 27, wherein the first wall part of the composite skin panel comprises a side edge joggle provided at a side edge of the composite skin panel, the side edge joggle permitting the first wall part of the frame element to overlap the composite skin panel's first wall part while maintaining a flush outer surface of the fuselage skin.

30. The fuselage structure according to claim 27, wherein the connection between a composite skin panel and another composite skin panel comprises joined second wall parts of both, wherein the second wall part of one composite skin panel comprises a joggle adjacent to the stiffener, the joggle permitting the second wall part of the composite skin panel to overlap with the other composite skin panel's second wall part while maintaining a flush outer surface of the fuselage skin.

31. The fuselage structure according to claim 27, wherein all of the first and/or all of the second wall parts are joined through an induction welded connection.

32. The fuselage structure according to claim 27, wherein the stiffeners of composite skin panels are connected through an induction welded connection to each other to form a continuous stringer.

33. The fuselage structure according to claim 27, wherein the fiber-reinforced composite skin panels and/or the frame elements are made of a fiber-reinforced composite material having a thermoplastic matrix.

34. An aircraft comprising the fuselage structure according to claim 27.

35. A method for manufacturing a fuselage structure of an aircraft, comprising: providing a plurality of frame elements spaced apart from one another in a direction parallel to a longitudinal axis of the aircraft and extending in a circumferential direction of the aircraft; providing a plurality of fiber-reinforced composite skin panels to extend between each pair of frame elements, such that a stiffener integrally formed in each composite skin panel extends radially inwards from an inner surface of each composite skin panel and in a direction parallel to the aircraft longitudinal axis; interconnecting the plurality of fiber-reinforced composite skin panels by joining second wall parts of composite skin panels; and to form an integrated fuselage skin part between each pair of frame elements; connecting the plurality of fiber-reinforced composite skin panels to each pair of frame elements by joining first wall parts of each composite skin panel to first wall parts of each frame element; and repeating the above steps for each pair of frame elements until a fuselage skin is formed, extending along a longitudinal axis of the aircraft and enclosing an inner space thereof, and being supported by the plurality of frame elements, wherein the first wall parts of the composite skin panels are located more radially inwards than the first wall parts of the frame element to which they are joined, wherein at least some of the first wall parts of the composite skin panels and the frame elements and/or at least some of the second wall parts of the composite skin panels are joined through an induction welded connection.

36. The method according to claim 35, wherein the frame elements have an I-shaped or H-shaped cross-section, and the first wall part of the frame element comprises a flange of the I- or H-shaped frame element.

37. The method according to claim 35, wherein the first wall part of the composite skin panel comprises a side edge joggle provided at a side edge of the composite skin panel, and the first wall part of the frame element is brought in overlapping arrangement with the composite skin panel's first wall part to permit maintaining a flush outer surface of the fuselage skin.

38. The method according to claim 35, wherein a composite skin panel is connected to another composite skin panel by joining second wall parts of both, wherein the second wall part of a composite skin panel comprises a side edge joggle provided at a side edge of the composite skin panel, and the second wall part of the other composite skin panel is brought in overlapping arrangement with the composite skin panel's second wall part to permit maintaining a flush outer surface of the fuselage skin.

39. The method according to claim 35, wherein all of the first and/or all of the second wall parts are joined through an induction welded connection.

40. The method according to claim 35, wherein the stiffeners of composite skin panels are connected to each other through an induction welded connection to form a continuous stringer.

41. The method according to claim 39, wherein the first and/or second wall parts are made of a fiber-reinforced composite material having a thermoplastic polymer matrix, and the thermally activated coupling means comprises the thermoplastic polymer matrix, and wherein joining the first and/or second wall parts is achieved by a method comprising pressurizing contact surfaces of the first and/or second wall parts to be joined, moving an inductor along the pressurized contacted surfaces of the first and/or second wall parts, generating an electromagnetic field in an induction-sensitive component, selected from carbon fibers, a metal or metal mesh, ferromagnetic particles, or combinations of these, of the first and/or second wall parts to heat the thermoplastic polymer of the first and/or second wall parts to above a melting temperature of the thermoplastic polymer, and connecting the contact surfaces of the first and/or second wall parts to each other by the molten thermoplastic polymer.

Description

BRIEF DESCRIPTION OF THE FIGURES

[0053] The invention will now be elucidated with reference to the following figures, without however being limited thereto. In the figures:

[0054] FIG. 1A schematically shows a perspective view of a fuselage structure according to the state of the art:

[0055] FIG. 1B schematically shows a cross-sectional view of the fuselage structure of FIG. 1A wherein the cross-section is taken in a plane extending parallel to a longitudinal axis of the fuselage structure;

[0056] FIG. 1C schematically shows a cross-sectional view of the fuselage of FIG. 1A wherein the cross-section is taken in a plane extending perpendicular to a longitudinal axis of the fuselage structure;

[0057] FIG. 2 schematically shows a perspective view of a fiber-reinforced composite skin panel in accordance with an embodiment of the invention;

[0058] FIG. 3 schematically shows a perspective view of a fiber-reinforced composite skin panel that is joined to another fiber-reinforced composite skin panel in accordance with an embodiment of the invention;

[0059] FIG. 4 schematically shows a perspective view of two fiber-reinforced composite skin panels that are joined to a frame element in accordance with an embodiment of the invention;

[0060] FIG. 5 schematically shows a cross-sectional view of an H-shaped frame element in accordance with an embodiment of the invention;

[0061] FIG. 6 schematically shows a perspective view of part of a fuselage structure in accordance with an embodiment of the invention;

[0062] FIG. 7 schematically shows a side view of a device for manufacturing the fiber-reinforced composite skin panels in accordance with an embodiment of the invention;

[0063] FIG. 8 schematically shows a cross-sectional side view of a tooling used in electromagnetic welding of fiber-reinforced composite skin panels to each other or to a frame element in accordance with an embodiment of the invention;

[0064] FIG. 9 schematically shows several stiffener designs in accordance with embodiments of the invention; and finally

[0065] FIG. 10 schematically shows a cross-section of two fiber-reinforced composite skin panels that are joined to a frame element in accordance with yet another embodiment of the invention.

DESCRIPTION OF EXEMPLARY EMBODIMENTS

[0066] Referring to FIG. 1, this disclosure relates to composite aircraft structures and more particularly to an aircraft fuselage structure 1 with a fuselage skin 2 made of aluminum or composite material. In the context of the present disclosure, the term “composite material” is intended to encompass fiber-reinforced composite materials (e.g., polymers) and advanced composite materials also known as advanced polymer matrix composites which generally comprise high strength fibers bound together by a matrix material or any known or other composite material(s) suitable for use in aircraft structural parts such as fuselage skins. Such composite materials may include fiber reinforcement materials such as carbon, aramid and/or glass fibers embedded into a thermosetting or thermoplastic matrix material. An exemplary fuselage structure 1 may be an aft fuselage section of an aircraft or any other portion of a fuselage structure 1 such as a cockpit section of fuselage for instance. The fuselage structure 1 may be a pressurized or non-pressurized section of fuselage. Fuselage structure 1 further comprises frame subassembly 3 comprising a plurality of frame elements 30. The frame elements 30 are preferably made of a fiber-reinforced composite material with a thermoplastic polymeric matrix, or may be made of a metallic material such as an aluminum-based alloy, a titanium-based alloy, steel or other suitable metallic material. When induction welding is used for connecting the composite skin panels 6 to metallic frame elements 30, the metal surfaces to be bonded or connected may require special surface preparation to ensure proper adhesion. The frame elements 30 may be interconnected though one or more longerons 4 to form the pre-assembled frame subassembly 3. The longerons 4 extend parallel to a longitudinal axis 11 of the fuselage structure 1. The frame elements 30 extend in a transverse plane 10 of the fuselage structure 1, which transverse plane 10 is perpendicular to the longitudinal axis 11 of the fuselage structure 1. Alternatively or in addition, one of more frame elements 30 may be positioned under an oblique angle relative to the longitudinal axis 11 of the fuselage structure. The frame elements 30 serve to provide support to an interior of the fuselage skin 2 and may be spaced apart along the longitudinal axis 11 of the fuselage structure 1. The relative spacing 31 of the frame elements 30 may be selected prior to assembly with the composite fuselage skin 2.

[0067] As also shown in FIG. 1, the fuselage structure 1 further comprises a number of stringers 5 that extend parallel to the longitudinal axis 11 of the fuselage structure 1 and penetrate the frame elements 30. The stringers 5 are generally less stiff as the longerons 4 and a number of these stringers 5 is provided between a pair of longerons 4. The purpose of the stringers 5 is to provide some shape stability, stiffness and buckling resistance to the fuselage skin 2 in between the longerons 4. As shown, the fuselage skin 2 is generally provided onto the structure consisting of frame elements 30, longerons 4 and stringers 5, and thereby provides a relatively smooth outer surface.

[0068] A fuselage skin 2 according to the state of the art may have a full barrel construction, which means that the fuselage skin 2 comprises a single piece that is closed in its circumferential direction and that extends completely around the longitudinal axis 11 of the aircraft. Such a composite skin 2 is typically manufactured using well known composite manufacturing methods such as automated fiber placement (AFP) or automated tape laying (ATL). Techniques such as ATL and AFP require large molds and use computer-guided robotics to lay one or several layers of reinforcing fiber tape or tows onto a mold or mandrel to form a part or structure. After tape-laying, the whole tape-laid structure is provided in an autoclave in order to harden the thermosetting matrix polymer of the tapes or tows. Autoclaving is typically performed under vacuum which requires wrapping the complete tape-laid structure in an air-impermeable foil. It goes without saying that autoclaving a complete tape-laid fuselage structure requirements large investments and is rather inefficient from an energetic and material waste point of view. However, no other manufacturing method has proven to be able to match the reliability of an autoclaved fuselage structure with a full barrel fuselage skin construction.

[0069] Referring now to FIG. 2, a fiber-reinforced composite skin panel 6 according to an embodiment of the present invention is shown. The panel 6 has two first side edges 60 that are each configured to be joined to a frame element 30 along a length 61 of said first side edges 60. A panel 6 thus extends between a pair of frame elements 30 and is connected thereto by a method that will be elucidated further below. The panel 6 further has two second side edges 62 that, in assembled state, extend in a direction parallel to the longitudinal axis 11 of the fuselage structure 1. An inner surface 63 of the panel 6 faces the inside of the assembled fuselage structure 1, while an outer surface 64 of the panel 6 forms part of an outer surface of the assembled fuselage skin 2. In order to accommodate the more or less cylindrical shape of an assembled fuselage skin 2, each composite skin panel 6 is curved around the longitudinal axis 11 with a radius of curvature R that may depend on the size of the aircraft and the number of panels 6 used to assemble the complete fuselage skin 2. The radius of curvature R may vary within a wide range, such as between 2 and 6 m. The second side edges 62 may be curved but preferably are substantially flat. The length 65 of a second side edge 62 also depends on the size of the aircraft and the relative spacing 31 of the frame elements 30.

[0070] The composite skin panel 6 of the invention further comprises a stiffener 66 integrally formed in each composite skin panel 6 and extending radially inwards from the inner surface 63, i.e. towards the longitudinal axis 11 of the fuselage structure 1. The stiffener 66 in the embodiment shown is hat-shaped with two upstanding walls 66-1 and a roof part 66-2. The stiffener 66 further extends along a line about parallel to the second side edges 62 of the panel 6 and to the aircraft longitudinal axis 11. The stiffener 66 in the shown embodiment has a hat-shaped cross-section. However this cross-section may have other shapes such as a H- or I-shape. For ease of manufacturing, a hat-shaped stiffener cross-section is preferred.

[0071] A way of providing the connection between a composite skin panel 6 and another composite skin panel 6 comprises joined second wall parts (67-1, 67-2) of both panels. The second wall parts (67-1, 67-2) extend about parallel to the second edges 62 of each panel 6 and comprise second wall parts 67-1 that are located in the vicinity or adjacent to the stiffener 66, and second wall parts 67-2 that are located at another side edge 62 of the panel 6, as shown in FIG. 2 for instance. In order to join two panels, the wall parts 67-1 of on panel 6 are brought in overlapping arrangement with the wall parts 67-2 of another panel 6, as shown in the top part of FIG. 3. In order for maintaining a flush outer surface 64 of the assembled fuselage skin 2, a second wall part 67-1 of one composite skin panel 6 comprises a joggle 68 adjacent to the stiffener 66. The joggle 68 extends along the second side edge 62 and permits a second wall part 67-2 of another composite skin panel 6 to overlap with the second wall part 67-1 and yet maintain a flush outer surface 64 between the two panels 6. The cooperating second wall parts (67-1, 67-2) are brought against each other and joined by any means, such as by mechanical bonding, adhesive bonding or induction welding bonding, the latter being preferred as will be elucidated further below.

[0072] First wall parts 69-1 and 69-2 extend about parallel to the first edges 60 of each panel 6 and comprise first wall parts 69-1 that are located at one first side edge 60, and first wall parts 69-2 that are located at an opposite first side edge 60 of the panel 6, as shown in FIG. 2 for instance. The first wall parts (69-1, 69-2) are configured to be joined to first wall parts (30-1, 30-2) of a frame element 30, as shown in FIG. 4. In the embodiment shown in FIGS. 4 and 5, the frame element 30 has an I-shaped cross-section and comprises two U-segments, positioned with their back webs (31-1, 31-2) against each other to form an I-shaped frame element 30, comprising a web plate (31-1, 31-2) and two flanges (30-1, 30-2) and (32-1, 32-2). Each flange may be covered with an additional flange plate (33-1, 33-2), if required. The flange plate 33-1 is positioned closer to the outside of the fuselage skin 2, while the flange plate 33-2 is positioned such that it points towards the inside of the fuselage skin 2 and the longitudinal axis 11. The different components (30, 31, 32) of a frame element 30 may be connected to each other by any means such as by mechanical bonding, adhesive bonding or induction welding bonding, the latter being preferred.

[0073] As shown in FIG. 4, the first wall parts (69-1, 69-2) of a composite skin panel 6 are joined to the first wall parts (30-1, 30-2) of the outer flange of the frame element 30. This preferably done such that the first wall parts (69-1, 69-2) of the composite skin panel 6 are located more radially inwards than the first wall parts (30-1, 30-2) of the frame element 30 to which it is joined. This may help in preventing failure between the first wall parts (69-1, 69-2) and (30-1, 30-2). Indeed, a pressurized fuselage will exert a pressure P from the inside to the outside onto the fuselage skin 2 (and fuselage skin panels 6) as shown in FIG. 4. This pressure P will tend to compress the first wall parts (69-1, 69-2) of a panel 6 and the first wall parts (30-1, 30-2) of a frame element.

[0074] The first wall parts (69-1, 69-2) of the composite skin panel 6 preferably also comprise a side edge joggle 70 provided at a side edge 60 of the composite skin panel 6. This side edge joggle 70 permits the first wall parts (30-1, 30-2) of the frame element (30) to overlap the composite skin panel's first wall parts (69-1, 69-2) while maintaining a flush outer surface 64 of the fuselage skin 2.

[0075] With reference to FIG. 6, a part 2a of a fuselage skin 2 according to an embodiment of the invention is shown. The skin part 2a is shown from the inside and comprises a curved frame element 30 having a radius of curvature R that corresponds to the radii of curvature R of each composite skin panel 6. A number of such panels 6 is connected to the frame element 3 along first edges 62, as has been described above. For clarity, the opposite first edges 62 are shown to be not connected to a frame element 30. However, to build a fuselage skin 2 these opposite first edges 62 are in reality connected to other frame elements 30, spaced at a distance 31 from an adjacent frame element 30. As also shown, the number of panels 6 that may be arranged between frame elements 30 may vary, for instance by adopting different panel lengths 61 in a direction transverse to the longitudinal axis 11 of the fuselage structure 1. In the embodiment shown, the panels 6 below the frame element 30 are twice as small as the panels 6 above said frame element 30.

[0076] In the assembled fuselage part 2a, the stiffeners 66 of the composite skin panels 6 are aligned with one another in a direction parallel to the fuselage structure 1 longitudinal axis 11.

[0077] Referring to FIG. 9, several embodiments of stiffeners 66 are schematically shown, and in particular the way in which they may be connected to other parts of the fuselage structure 1, such as frame elements 30. In conventional fuselage structures 1, the stringers 5 generally run along the complete length of the fuselage structure 1 and extend trough the frame elements 30. According to the invention in which the panels 6 comprise integrally formed stiffeners 66, it is inevitable that these stiffeners 66 may not be much longer than the distance 31 (or pitch length) between two frame elements 30. The intermittent stiffeners 66 may be aligned to form a pseudo-stringer structure but may provide a less favorable buckling resistance to the fuselage skin 2 than the conventional continuous stringer 5. To at least partly overcome this disadvantage, a number of stiffener 66 designs are proposed. In the stiffener designs shown in FIGS. 9C and 9E, each stiffener 66 is connected to a frame element 30, either by a separate connector 90, or by providing the stiffener 66 with connector lips 91, as shown. In the designs shown in FIGS. 9A and 9B, the stiffener 66 does not directly connects to the frame element 30, but rather through the first wall parts (69-1, 69-2) of the panel 6.

[0078] Besides the composite skin panels 6, the frame elements 30 may also be made of a fiber-reinforced composite material having a thermoplastic matrix, such as a carbon/PEEK composite material for instance.

[0079] As shown in the cross-sectional view of FIG. 10, the first wall parts (69-1, 69-2) of a composite skin panel 6 are joined to the first wall parts (30-1, 30-2) of the outer flange of the frame element 30. This is preferably done such that the first wall parts (69-1, 69-2) of the composite skin panel 6 are located more radially inwards than the first wall parts (30-1, 30-2) of the frame element 30 to which it is joined, for the reasons given above. As with the embodiment shown in FIG. 4, the first wall parts (69-1, 69-2) of the composite skin panel 6 preferably also comprise a side edge joggle 70 provided at a side edge 60 of the composite skin panel 6. This is to allow the first wall parts (30-1, 30-2) of the frame element (30) to overlap the composite skin panel's first wall parts (69-1, 69-2) while maintaining a flush outer surface 64 of the fuselage skin 2. In the embodiment shown, an outer surface of (part of) the frame element 30 and of each panel 6 is provided with a conducting structure or material, such as a metal mesh (30-3, 6-3). This allows the connection between the first wall parts (69-1, 69-2) of each composite skin panel 6 and the first wall parts (30-1, 30-2) of the outer flange of the frame element 30 to be made by induction welding. At the same token, the outer surfaces, comprising the conducting structure or material in the form of a metal mesh (30-3, 6-3), are made electrically conducting, which provides protection against lightning strike.

[0080] The fiber-reinforced composite skin panels 6 may conveniently be made by press forming. Referring to FIG. 7, a number of carbon fiber unidirectional tapes having a thermoplastic polymer matrix such as PEEK is cut to size and stacked into a desired layup 7, such as a [+45°/−45°/0°/90°].sub.s layup 7. The layup 7 may be sandwiched between two metal plates (not shown) and a vacuum is applied between the plates to reduce trapped air, while the atmospheric pressure compresses the layup 7 somewhat. The layup 7 is then transferred into an oven 8 in which the thermoplastic polymer matrix is allowed to melt, which enhances flow of matrix polymer between the tapes and helps in fusing the same. The layup 7′ is transferred to a pre-heated mold 9 provided at a temperature below the softening temperature T.sub.s of the thermoplastic matrix polymer. The mold 9 comprises a male mold part 9-1 and a shape-matched female mold part 9-2, both of which mold parts (9-1, 9-2) are brought together and pressed against each other with a pre-determined force 50, which is held until the laminate is again cooled to a temperature below the glass transition temperature T.sub.g of the matrix polymer of the produced panel 6 by conduction heat to the cooler molds. The shape of the panel 6 is now consolidated and the edges of the panel 6 may be trimmed to size. The need for only one pair of relatively small mold parts (9-1, 9-2) may significantly reduce production and investment costs. Suitable panels 6 may be mostly single-curved, which reduces the risk of having misaligned fibers in the panels 6 as produced.

[0081] A method for manufacturing a fuselage structure 1 of an aircraft using the invention provides a plurality of frame elements 30 spaced apart from one another over a distance 31 in a direction parallel to the longitudinal axis 11 of the aircraft or fuselage structure 1. The frame elements 30 each extend in a plane 10 that is about perpendicular to the longitudinal axis 11. A plurality of fiber-reinforced composite skin panels 6 is provided between each pair of frame elements 30. The panels 6 are oriented such that a stiffener 66 formed in each composite skin panel 6 extends radially inwards from an inner surface 63 of each composite skin panel 6 and in a direction parallel to the aircraft longitudinal axis 11, as for instance shown in FIG. 6. The plurality of fiber-reinforced composite skin panels 6 is then interconnected by joining second wall parts (67-1, 67-2) of each composite skin panel 6 to second wall parts (67-1, 67-2) of another composite skin panel 6. In this manner, an integrated fuselage skin part 2a may be formed between each pair of frame elements 30. The process also requires connecting the plurality of fiber-reinforced composite skin panels 6 to each pair of frame elements 30 by joining first wall parts (69-1, 69-2) of each composite skin panel 6 with first wall parts (30-1, 30-2) of each frame element 30. The above steps may be repeated for each pair of frame elements 30 until a fuselage skin 2 is formed, extending along the longitudinal axis 11 of the fuselage structure 1 and enclosing an inner space thereof.

[0082] A preferred manner to connect the fiber-reinforced composite skin panels 6 to each other and/or to each pair of frame elements 30 comprises electromagnetic welding, as schematically shown in FIG. 8. The molded parts (40, 41) shown in FIG. 8 may comprise two fiber-reinforced composite skin panels 6, a fiber-reinforced composite skin panel 6 and a frame element 30, or two stiffeners 66.

[0083] The molded parts (40, 41) preferably are to be connected by electromagnetic welding. As referred to elsewhere, both molded parts (40, 41) are preferably manufactured from a thermoplastic matrix polymer reinforced with carbon fibres, wherein the carbon fibres also serve as induction-sensitive component for heating of the thermoplastic polymer matrix for the purpose of welding. The molded parts (40, 41) need to be joined along molded wall parts (40a, 41a) that are brought together to define a common contact surface 42 for coupling. An inductor 43, which may for instance be a linear inductor 43 that provided a substantial cylindrical electromagnetic field heats the molded wall parts (40a, 41a) and their common contact surface 42 to a temperature which is high enough to thermally activate the thermoplastic matrix polymer, or, optionally, a thermally activated adhesive applied to the contact surface 42. The inductor may be moved along the wall parts (40a, 41a) without making physical contact with the wall parts (40a, 41a). During heating and/or optionally a short time thereafter, the thermally activated contact surface 42 is compressed in the direction A by appropriate tooling, comprising a clamping tool 44 and a static plate 45 provided on top of the wall parts (40a, 41a). A well developed and strong connection between the molded parts (40, 41) may be made in this way.

[0084] The invention is not limited to the above given examples and variations thereto may be envisaged within the scope of the appended claims.