# US 11,440,283 B2 · Composite sandwich panels with over-crushed edge regions

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- Number: US 11,440,283 B2
- Document ID: US11440283B2
- Status: Granted
- Kind: B2
- Application number: 15/887,839
- Filed: 2018-02-02
- Published: 2019-08-08
- Granted: 2022-09-13
- Earliest claimed: 2018-02-02
- Assignees: The Boeing Company
- Applicants: The Boeing Company
- Inventors: Jason Turner; Mitchell Zachary Kotlik; Christopher Robert Schwartz; Michael B. Black
- Web page: https://patentgrep.com/patent/US11440283B2

## CPC

- B29C 43/02: of articles of definite length, i.e. discrete articles {(B29C35/0227 takes precedence)}
- B29C 43/203: {Making multilayered articles}
- B29C 43/36: Moulds for making articles of definite length, i.e. discrete articles
- B29C 43/40: with means for cutting the article
- B29C 70/026: {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)}
- B29C 70/086: {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)}
- B29C 70/46: using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs
- B29C 70/545: {Perforating, cutting or machining during or after moulding}
- B29C 70/763: {the edges being disposed in a substantial flat plane}
- B29C 70/865: {completely encapsulated}
- B29D 99/0014: {provided with ridges or ribs, e.g. joined ribs}
- B29D 99/0089: {Producing honeycomb structures (consisting of porous ceramic ware C04B38/0006)}
- B32B 3/02: characterised by features of form at particular places, e.g. in edge regions {(non-uniform thickness B32B3/263)}
- B32B 3/12: characterised by a layer of regularly- arranged cells, e.g. a honeycomb structure
- B32B 3/30: characterised by a layer formed with recesses or projections, e.g. {hollows, grooves, protuberances, ribs (apertured layer B32B3/266; layer with cavities or internal voids B32B3/26)}
- B29C 70/685: {by laminating inserts between two plastic films or plates}
- B29C 70/72: Encapsulating inserts having non-encapsulated projections, e.g. extremities or terminal portions of electrical components {(B29C70/742 takes precedence)}
- B29C 2043/403: {knife blades}
- B29C 2793/0027: Cutting off
- B29C 2793/009: after shaping
- B29L 2031/3076: {Aircrafts (blades, propellers B29L2031/08)}
- B29L 2031/608: {Honeycomb structures}
- B29L 2031/72: {Cutting equipment, e.g. punches (sawing equipment B29L2031/77; grinding equipment B29L2031/736)}
- B32B 2250/40: Symmetrical or sandwich layers, e.g. ABA, ABCBA, ABCCBA
- B32B 2605/18: Aircraft
- B29C: SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING (making preforms B29B11/00; making laminated products by combining previously unconnected layers which become one product whose layers will remain together B32B37/00 - B32B41/00)
- B29D: PRODUCING PARTICULAR ARTICLES FROM PLASTICS OR FROM SUBSTANCES IN A PLASTIC STATE (making granules B29B9/00; making preforms B29B11/00)
- B32B: LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B29L: INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES

## 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.

## Claims

### Claim 1

A method comprising:
loading a first composite skin, a second composite skin, and a hollow cell core into a tool comprising opposing dies;
closing the tool to form an interior shape of the tool between the opposing dies, the interior shape comprising a periphery region; and
compressing the first composite skin, the second composite skin, and the hollow cell core using the opposing dies to form and consolidate a composite sandwich panel having a nominal thickness while applying an over-crush pressure to portions of the composite sandwich panel to form a first over-crush edge region comprising an edge and a transition region, the edge having a first thickness at least 40% less than the nominal thickness of the composite sandwich panel, the nominal thickness being the thickness of the composite sandwich panel away from the first over- crush edge region after consolidating the first composite skin, the second composite skin, and the hollow cell core, wherein the periphery region surrounds the first over-crush edge region when the composite sandwich panel is in the tool and provides a path for volatiles, released from the composite sandwich panel during compression, to escape the tool between the opposing dies.

### Claim 2 (depends on 1)

The method of claim 1, wherein the periphery region has a thickness greater than the nominal thickness.

### Claim 3 (depends on 1)

The method of claim 1 further comprising:
reducing a thickness of the composite sandwich panel from the nominal thickness to the first thickness in the transition region using an over-crush forming feature of one of the opposing dies, wherein a length of the transition region is at least 0.25 inches.

### Claim 4 (depends on 3)

The method of claim 3, wherein the over-crush forming feature comprises a protrusion.

### Claim 5 (depends on 1)

The method of claim 1 further comprising:
cutting the first over-crush edge region using numerically controlled machinery to create a cut edge having a flat appearance without visible core cells.

### Claim 6 (depends on 5)

The method of claim 5, wherein cutting the first over-crush edge region removes a portion of composite sandwich panel comprising the first thickness.

### Claim 7 (depends on 1)

The method of claim 1, wherein the periphery region is substantially perpendicular to a direction the opposing dies are moveable relative to each other as the periphery region extends away from the first over-crush edge region towards an outside of the tool.

### Claim 8 (depends on 1)

The method of claim 1 further comprising:
forming a second over-crush edge region and a third over-crush edge region during compressing of the first composite skin, the second composite skin, and the hollow cell core; and
allowing volatiles to escape through a fourth edge of the composite sandwich panel while forming the second over-crush edge region and the third over-crush edge region of the composite sandwich panel.

### Claim 9 (depends on 8)

The method of claim 8, wherein the periphery region surrounds the second over-crush region, the third over-crush edge region, and the fourth edge when the composite sandwich panel is in the tool.

### Claim 10 (depends on 1)

The method of claim 1, wherein the edge has a flat appearance with no visible core cells.

### Claim 11 (depends on 1)

The method of claim 1, wherein the edge is a rolled edge.

### Claim 12 (depends on 1)

The method of claim 1, further comprising paint covering the edge.

### Claim 13 (depends on 1)

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

### Claim 14 (depends on 1)

The method of claim 1, wherein the composite sandwich panel further comprises an open edge having visible core cells, wherein the open edge is the nominal thickness of the composite sandwich panel.

### Claim 15 (depends on 14)

The method of claim 14, wherein the composite sandwich panel further comprises an edge treatment applied to the open edge.

### Claim 16 (depends on 1)

The method of claim 1, wherein the periphery region has a thickness at least twice the nominal thickness.

### Claim 17 (depends on 1)

The method of claim 1, wherein the first thickness is in a range of 0.050 inches to 0.10 inches.