US 11,939,038 B2
Fuselage structure of an aircraft and method for manufacturing the same
Maarten Labordus, Delft (NL)
Assigned to KOK & VAN ENGELEN COMPOSITE STRUCTURES B.V., The Hague (NL)
Appl. No. 17/625,203
Filed by KOK & VAN ENGELEN COMPOSITE STRUCTURES B.V., The Hague (NL)
PCT Filed Jun. 25, 2020, PCT No. PCT/NL2020/050419
§ 371(c)(1), (2) Date Jan. 6, 2022,
PCT Pub. No. WO2021/006725, PCT Pub. Date Jan. 14, 2021.
Claims priority of application No. 2023459 (NL), filed on Jul. 8, 2019.
Prior Publication US 2022/0258847 A1, Aug. 18, 2022
Int. Cl. B64C 1/12 (2006.01); B29C 65/36 (2006.01); B29D 99/00 (2010.01); B29L 31/30 (2006.01); B64C 1/06 (2006.01); B64C 1/00 (2006.01)
CPC B64C 1/12 (2013.01) [B29C 65/3636 (2013.01); B29C 65/3684 (2013.01); B29D 99/0014 (2013.01); B64C 1/061 (2013.01); B64C 1/064 (2013.01); B64C 1/069 (2013.01); B29L 2031/3082 (2013.01); B64C 2001/0072 (2013.01)] 13 Claims
OG exemplary drawing
 
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.