Composite mandrel for autoclave curing applications

Patent metadata

Field	Value
Publication number	US-9327467-B2
Application number	US-17084308-A
Country	US
Kind code	B2
Filing date	Jul 10, 2008
Priority date	Jul 10, 2008
Publication date	May 3, 2016
Grant date	May 3, 2016

Abstract

A composite mandrel includes a generally elongated mandrel body comprising a resilient mandrel core and an elastomeric mandrel outer layer disposed outside the mandrel core. A method for fabricating a contoured stiffened composite panel is also disclosed.

First claim

What is claimed is:

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.

2. The method of claim 1 wherein said tooling surface comprises a generally concave contour.

3. The method of claim 1 wherein the one-piece resilient mandrel body comprises a generally triangular cross-section.

4. The method of claim 1 wherein the one-piece resilient mandrel body comprises a generally trapezoidal cross-section.

5. The method of claim 1 wherein the elastomeric outer layer comprises rubber.

6. The method of claim 1 , wherein removing includes deforming the elastomeric mandrel outer layer.

7. A method for fabricating a contoured stiffened composite panel for an aircraft structure, comprising:
providing a tooling surface having a generally concave contour;
placing a base composite layer on said tooling surface;
placing at least one stiffening element having a stiffening element cavity on said base composite layer, the stiffening element comprising at least a stiffening element base surface and a stiffening element side surface;
providing a composite mandrel, said composite mandrel comprising a generally triangular or trapezoidal cross-section and including a resilient foam mandrel one-piece core coextensive with an elastic rubber mandrel outer layer disposed outside said resilient foam mandrel one-piece core, the composite mandrel comprising a composite mandrel base surface and a composite mandrel side surface, wherein a cross-sectional area and type of foam used for the resilient foam mandrel one-piece 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 during curing, 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 forming the elastic rubber mandrel outer layer to have a substantially constant thickness;
inserting said composite mandrel in said stiffening element cavity of said at least one stiffening element such that the composite mandrel base surface contacts the stiffening element base surface and the composite mandrel side surface contacts the stiffening element side surface;
enclosing said base composite layer and said at least one stiffening element in a vacuum bag and curing, during curing the composite mandrel base surface maintaining contact with the stiffening element base surface and the composite mandrel side surface maintaining contact with the stiffening element side surface so as to prevent collapse of the stiffening element; and
removing said composite mandrel, after curing and without further heating, from said stiffening element cavity of said at least one stiffening element, the removing including deforming the elastic rubber mandrel outer layer so as to reduce an effort of removal.

8. The method of claim 7 wherein the composite mandrel comprises a generally triangular cross-section.

9. The method of claim 7 wherein the composite mandrel comprises a generally trapezoidal cross-section.

10. A method for fabricating a composite panel with a stringer, comprising:
placing a base composite layer on a tooling surface, the tooling surface having a generally concave contour;
placing a stiffening element having a cavity on the base composite layer, the stiffening element comprising at least a stiffening element base surface and a stiffening element side surface;
inserting a composite mandrel in the cavity of the stiffening element, the composite mandrel comprising a core comprising a foam and an outer layer comprising an elastomeric material, the composite mandrel comprising a composite mandrel base surface and a composite mandrel side surface, the composite mandrel substantially filling the cavity of the stiffening element such that the composite mandrel base surface contacts the stiffening element base surface and the composite mandrel side surface contacts the stiffening element side surface, wherein a cross-sectional area and type of foam used for the core is engineered to impart compression compliance under autoclave pressure to offset a combined thermal expansion behavior of the core and the outer layer, wherein in being engineered, the core provides structural and compressive support necessary to maintain a shape of the stringer during automated composite fiber placement as well as autoclave curing, and wherein being engineered further comprises forming the outer layer to have a substantially constant thickness;
enclosing the base composite layer and the stiffening element in a vacuum bag;
curing the base composite layer and the stiffening element so as to form the composite panel with the stringer during curing the composite mandrel substantially filling the cavity of the stiffening element such that the composite mandrel base surface contacts the stiffening element base surface and the composite mandrel side surface contacts the stiffening element side surface; and
removing the unitary composite mandrel from the cavity of the stiffening element, the removing including deforming the elastomeric material of the outer layer so as to reduce an effort associated with the removing.

11. The method of claim 10 wherein the outer layer of the composite mandrel comprises an elastic rubber mandrel outer layer.

12. The method of claim 10 , wherein the composite mandrel comprises a generally triangular cross-section.

13. The method of claim 10 , wherein the composite mandrel comprises a generally trapezoidal cross-section.

Assignees

• Robins Brian G
• Rotter Daniel M
• Washburn Todd J
• George Panagiotis E
• Boeing Co

Inventors

• Robins Brian G
• Rotter Daniel M
• Washburn Todd J
• George Panagiotis E

Classifications

• B29D99/0014 Primary

  provided with ridges or ribs, e.g. joined ribs · CPC title

• B29C33/505 Primary

  cores or mandrels, e.g. inflatable (B29C33/0016 takes precedence; for winding and joining B29C53/824; for supporting articles during joining B29C66/634; flexible cores for vulcanizing tyres B29D30/0654) · CPC title

• B29C33/3814 Primary

  Porous moulds (adapted for vacuum forming B29C51/365) · CPC title

• B29C70/865

  completely encapsulated · CPC title

• B29L2031/3076

  Aircrafts (blades, propellers B29L2031/08) · CPC title