Methods and apparatus for curing composite nacelle structure

Patent metadata

Field	Value
Publication number	US-11400620-B2
Application number	US-201816191699-A
Country	US
Kind code	B2
Filing date	Nov 15, 2018
Priority date	Oct 1, 2014
Publication date	Aug 2, 2022
Grant date	Aug 2, 2022

Abstract

Methods and apparatus for curing curved cylinder-like workpieces (e.g., in the shape of a half or full barrel) made of composite material, such as nacelle honeycomb core composite sandwich structures. These methods enable tailored curing of composite nacelle structures, to significantly reduce capital cost and fabrication cycle time. In lieu of an autoclave or oven, a pressurized ring-shaped cure volume is defined by a partitioned enclosure that mimics the cylinder-like shape of the composite nacelle structure with only limited clearance (e.g., a partitioned enclosure comprising inner and outer concentric cylinder-like walls). A tool (e.g., a mandrel) and at least one composite nacelle structure supported thereon are placed in the cure volume for curing. Integrally heated tooling, optionally in combination with other heating methods, such as infrared heaters, is utilized to provide the temperature profile necessary for cure.

First claim

The invention claimed is:

1. A method for curing a composite structure, comprising: (a) forming a tool-composite structure assembly by placing an uncured composite structure in contact with an outer surface of a hollow tool having a closed contour; (b) placing the tool-composite structure assembly on a circular manifold which is attached to a base, the circular manifold being configured to couple the hollow tool to a source of energy; (c) enclosing a ring-shaped cure volume having an outer boundary that surrounds the tool-composite structure assembly with clearance from the uncured composite structure and a lower boundary formed in part by the base; (d) heating the uncured composite structure during a cure cycle while the uncured composite structure is positioned within the ring-shaped cure volume; and (e) producing a specified pressure inside the ring-shaped cure volume during the cure cycle.

2. The method as recited in claim 1 , further comprising: removing the tool-composite structure assembly from the ring-shaped cure volume upon completion of the cure cycle; and demolding the composite structure from the hollow tool.

3. The method as recited in claim 2 , wherein the composite structure is a component of an aircraft.

4. The method as recited in claim 3 , wherein the aircraft component is one of the following components: an inlet inner acoustic panel, a fan cowl panel, a thrust reverser outer acoustic panel, a thrust reverser outer cowl panel or a thrust reverser inner wall panel.

5. The method as recited in claim 1 , wherein the uncured composite structure has a closed contour, surrounds the hollow tool and is surrounded by the outer boundary of the ring-shaped cure volume.

6. The method as recited in claim 1 , further comprising shaping the surface of the hollow tool to conform to an inner mold line of the uncured composite structure prior to step (a).

7. The method as recited in claim 1 , wherein step (b) comprises indexing the hollow tool to the base.

8. The method as recited in claim 1 , wherein step (c) comprises lowering a sleeve to a position whereat a lower end of the sleeve is supported by the base and the sleeve surrounds the hollow tool, the sleeve forming an outer boundary of the ring-shaped cure volume.

9. The method as recited in claim 8 , wherein step (c) further comprises lowering a plug to a position whereat a lower end of the plug is supported by the base and the hollow tool surrounds the plug, the plug forming an inner boundary of the ring-shaped cure volume.

10. The method as recited in claim 8 , further comprising attaching or connecting the sleeve to a top plate prior to step (c), wherein step (c) comprises lowering the top plate and the sleeve to respective positions whereat a lower end of the sleeve is supported by the base and the sleeve surrounds the hollow tool, the sleeve forming an outer boundary of the ring-shaped cure volume, and the top plate forming an upper boundary of the ring-shaped cure volume.

11. A method for curing a composite structure, comprising: (a) forming a tool-composite structure assembly by placing an uncured composite structure in contact with an outer surface of a hollow tool having a closed contour and having integrated heating elements; (b) placing the tool-composite structure assembly on a circular manifold which is attached to a base, the circular manifold being configured to couple the integrated heating elements to a source of energy; (c) enclosing a ring-shaped cure volume having an outer boundary that surrounds the hollow tool with clearance from the uncured composite structure and a lower boundary formed in part by the base; (d) activating the integrated heating elements to heat the uncured composite structure during a cure cycle; and (e) producing a specified pressure inside the ring-shaped cure volume during the cure cycle.

12. The method as recited in claim 11 , further comprising: removing the tool-composite structure assembly from the ring-shaped cure volume; and demolding the composite structure from the hollow tool.

13. The method as recited in claim 11 , wherein the uncured composite structure has a closed contour, surrounds the hollow tool and is surrounded by the outer boundary of the ring-shaped cure volume.

14. A method for curing a composite structure, comprising: (a) placing an uncured composite structure in contact with an outer surface of a hollow tool having a closed contour; (b) placing the hollow tool on a circular manifold which is attached to a base, the circular manifold being configured to couple the tool to a source of energy; (c) assembling a top plate, the base, an outer wall having a closed contour and an inner wall having a closed contour to form an enclosure that bounds a ring-shaped cure volume which is partly occupied by and has an outer boundary not in contact with the uncured composite structure; (d) heating the uncured composite structure during a cure cycle while the uncured composite structure is positioned within the ring-shaped cure volume; and (e) producing a specified pressure inside the ring-shaped cure volume during the cure cycle.

15. The method as recited in claim 14 , further comprising integrating a multiplicity of heating elements in the hollow tool, wherein step (d) comprises supplying heat to the hollow tool via the circular manifold and the heating elements.

16. The method as recited in claim 15 , wherein the heating elements supply heat to the hollow tool by transforming electric current into heat.

17. The method as recited in claim 15 , wherein the heating elements supply heat to the hollow tool by carrying heated fluid.

18. The method as recited in claim 14 , further comprising: removing the top plate; separating the hollow tool from the base; and demolding the composite structure from the tool.

19. The method as recited in claim 15 , wherein the uncured composite structure has a closed contour, surrounds the hollow tool and is surrounded by the outer wall as a result of step (c).

20. The method as recited in claim 14 , wherein the composite structure is a component of an aircraft.

Assignees

• Boeing Co

Inventors

• Riedel Brian L

Classifications

• B29C2035/043

  oil · CPC title

• B29C2035/0211

  resistance heating (B29C2035/0811 takes precedence) · CPC title

• B29C2035/0822

  using IR radiation · CPC title

• B29C35/0227 Primary

  using pressure vessels, e.g. autoclaves, vulcanising pans (B29C35/065 takes precedence) · CPC title

• B29C35/04

  using liquids, gas or steam · CPC title