Fuselage structure of an aircraft and method for manufacturing the same

1. 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 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; and

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.

2. The fuselage structure according to claim 1 , 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.

3. The fuselage structure according to claim 1 , 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.

4. The fuselage structure according to claim 1 , wherein all of the first and/or all of the second wall parts are joined through an induction welded connection.

5. The fuselage structure according to claim 1 , wherein the stiffeners of composite skin panels are connected through an induction welded connection to each other to form a continuous stringer.

6. The fuselage structure according to claim 1 , wherein the fiber-reinforced composite skin panels and/or the frame elements are made of a fiber-reinforced composite material having a thermoplastic matrix.

7. An aircraft comprising the fuselage structure according to claim 1 .

8. 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 the 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, and

wherein further 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.

9. The method according to claim 8 , 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.

10. The method according to claim 8 , 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.

11. The method according to claim 8 , wherein all of the first and/or all of the second wall parts are joined through an induction welded connection.

12. The method according to claim 11 , wherein the first and/or second wall parts are made of a fiber-reinforced composite material having a 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.

13. The method according to claim 8 , wherein the stiffeners of composite skin panels are connected to each other through an induction welded connection to form a continuous stringer.