Composite sandwich panels with over-crushed edge regions

Patent metadata

Field	Value
Publication number	US-11926077-B2
Application number	US-202217817064-A
Country	US
Kind code	B2
Filing date	Aug 3, 2022
Priority date	Feb 2, 2018
Publication date	Mar 12, 2024
Grant date	Mar 12, 2024

Abstract

A composite sandwich panel comprises a first composite skin, a second composite skin, a hollow cell core between the first composite skin and the second composite skin, and a first over-crush edge region with a first edge. The first edge has a first thickness at least 40% less than a nominal thickness of the composite sandwich panel. The first over-crush edge region has a length of at least 0.25 inches over which a thickness of the composite sandwich panel decreases.

First claim

What is claimed is:

1. A composite sandwich panel comprising:
a first composite skin;
a second composite skin;
a hollow cell core between the first composite skin and the second composite skin, wherein the first composite skin and the second composite skin and the hollow cell core are consolidated resulting in a 5% to 10% reduction in thickness of the entire composite panel to a nominal thickness;
a first edge comprising a first over-crush edge region, the first over-crush edge region having a first thickness at least 40% less than the nominal thickness of the composite sandwich panel, wherein the first over-crush edge region has a length of at least 0.25 inches over which a thickness of the composite sandwich panel decreases; and
an open edge, wherein the open edge is the nominal thickness of the composite sandwich panel.

2. The composite sandwich panel of claim 1 , wherein the first edge has a flat appearance with no visible core cells.

3. The composite sandwich panel of claim 1 , wherein the first edge is a rolled edge.

4. The composite sandwich panel of claim 1 , further comprising paint covering the first edge.

5. The composite sandwich panel of claim 1 , wherein the composite sandwich panel forms a light valence of an aircraft, and wherein the first edge is visible to passengers on the aircraft.

6. The composite sandwich panel of claim 1 further comprising:
a second edge comprising a second over-crush edge region, the second over-crush edge region having a second thickness at least 40% less than the nominal thickness of the composite sandwich panel, wherein the second over-crush edge region has a length of at least 0.25 inches over which the thickness of the composite sandwich panel decreases; and
a third edge comprising a third over-crush edge region, the third over-crush edge region having a third thickness at least 40% less than the nominal thickness of the composite sandwich panel, wherein the third over-crush edge region has a length of at least 0.25 inches over which the thickness of the composite sandwich panel decreases.

7. The composite sandwich panel of claim 1 , wherein the open edge has visible core cells and wherein the open edge is not over-crushed and is configured to allow volatile compounds to escape during consolidation of the first composite skin, the second composite skin, and the hollow cell core.

8. The composite sandwich panel of claim 7 further comprising an edge treatment applied to the open edge.

9. A composite sandwich panel comprising:
a first composite skin;
a second composite skin;
a hollow cell core between the first composite skin and the second composite skin, wherein the first composite skin and the second composite and the hollow cell core are consolidated by a crush pressure applied to an entirety of the first composite skin and the second composite skin in contact with the hollow cell core resulting in a 5% to 10% reduction in thickness of the entire composite panel to a nominal thickness;
a first over-crush edge region formed by an over-crush pressure applied to portions of the first composite skin, the second composite skin, and the hollow cell core, the first over-crush edge region having a first thickness at least 40% less than the nominal thickness of the composite sandwich panel, the nominal thickness being a thickness of the composite sandwich panel away from the first over-crush edge region, wherein the nominal thickness is in a range of 0.083 inches to 0.50 inches, and wherein the first over-crush edge region has a length of at least 0.25 inches over which the thickness of the composite sandwich panel decreases; and
an open edge having no over-crush edge region, wherein the open edge is the nominal thickness of the composite sandwich panel.

10. The composite sandwich panel of claim 9 wherein the first over-crush edge region comprises an edge and a transition region, the edge having the first thickness and the transition region having the length of at least 0.25 inches over which the thickness of the composite sandwich panel decreases from the nominal thickness to the first thickness.

11. The composite sandwich panel of claim 10 , wherein the edge has a flat appearance with no visible core cells.

12. The composite sandwich panel of claim 10 , wherein the edge is a rolled edge.

13. The composite sandwich panel of claim 10 , further comprising paint covering the edge.

14. The composite sandwich panel of claim 10 , wherein the composite sandwich panel forms a light valence of an aircraft, and wherein the edge is visible to passengers on the aircraft.

15. The composite sandwich panel of claim 9 , wherein the first over-crush edge region is configured to be formed by an interior shape of two opposing dies under compression and surrounded by a periphery region of the interior shape.

16. The composite sandwich panel of claim 15 , wherein the periphery region has a thickness greater than the nominal thickness.

17. The composite sandwich panel of claim 16 , wherein during formation of the first over-crush edge region, the periphery region provides a path for volatiles, released from the composite sandwich panel during compression.

18. The composite sandwich panel of claim 9 , wherein the first over-crush edge region is formed by the over-crush pressure applied to the first composite skin, the second composite skin, and the hollow cell core while the first composite skin, the second composite skin, and the hollow cell core are compressed.

19. The composite sandwich panel of claim 9 , wherein the first over-crush edge region comprises a first edge and a first transition region, the composite sandwich panel further comprising:
a second over-crush edge region with a second edge, the second edge having a second thickness at least 40% less than the nominal thickness of the composite sandwich panel, wherein the second over-crush edge region has a length of at least 0.25 inches over which the thickness of the composite sandwich panel decreases; and
a third over-crush edge region with a third edge, the third edge having a third thickness at least 40% less than the nominal thickness of the composite sandwich panel, wherein the third over-crush edge region has a length of at least 0.25 inches over which the thickness of the composite sandwich panel decreases.

20. The composite sandwich panel of claim 19 , wherein the second over-crush edge region and the third over-crush edge region are formed by the over-crush pressure applied to portions of the first composite skin, the second composite skin, and the hollow cell core while the first composite skin, the second composite skin, and the hollow cell core are compressed.

Assignees

• Boeing Co

Inventors

• Turner Jason
• Kotlik Mitchell Zachary
• Schwartz Christopher Robert
• Black Michael B

Classifications

• B29C43/02 Primary

  of articles of definite length, i.e. discrete articles {(B29C35/0227 takes precedence)} · CPC title

• B29C43/203

  Making multilayered articles · CPC title

• B29C43/36

  Moulds for making articles of definite length, i.e. discrete articles · CPC title

• B29C43/40

  with means for cutting the article · CPC title

• B29C70/026

  and with one or more layers of pure plastics material, e.g. foam layers (applying a non-preformed coating, e.g. a gel-coat B29C37/0025; with foam blocks B29C70/86) · CPC title