Patent ID: 9327467
Filing Date: 2016-05-03
CPC Classification: B29C,B29D,B29L

Claim Text:
1. A method for fabricating a contoured stiffened composite panel for an aircraft structure, comprising: placing a base composite layer on a tooling surface; placing at least one stiffening element having a stiffening element cavity on said base composite layer; inserting a one-piece resilient mandrel body in said stiffening element cavity, the one-piece resilient mandrel body comprising a foam core and an elastomeric outer layer substantially co-extensive with the foam core, and wherein the one-piece resilient mandrel body substantially fills the cavity of the stiffening element; wherein a cross-sectional area and type of foam used for the foam core is engineered to impart compression compliance under autoclave pressure to offset a combined thermal expansion behavior of the foam core and the elastomeric outer layer, wherein in being engineered, the foam core provides structural and compressive support necessary to maintain a shape of the contoured stiffened composite panel during automated composite fiber placement as well as autoclave curing, and wherein being engineered further comprises the elastomeric outer layer having a substantially constant thickness; enclosing said base composite layer and said at least one stiffening element in a vacuum bag and curing the base composite layer and the at least one stiffening element; and removing the one-piece resilient mandrel body from said stiffening element cavity of said at least one stiffening element following said curing.