Integrally stiffened bonded panel with vented pockets and methods of manufacture

Patent metadata

Field	Value
Publication number	US-11338903-B2
Application number	US-202016882889-A
Country	US
Kind code	B2
Filing date	May 26, 2020
Priority date	Oct 21, 2019
Publication date	May 24, 2022
Grant date	May 24, 2022

Abstract

Methods, systems, and apparatuses are disclosed for the manufacture of composite components having incorporated reinforcing structures machined into composite material substrates, and composite components manufactured according to disclosed methods, and assemblies and larger structures comprising the composite material components.

First claim

What is claimed is:

1. A method comprising:
orienting a plurality of composite material layers onto a tool to form a composite material substrate wing assembly, said composite material substrate wing assembly comprising a composite material substrate thickness, said tool comprising a tooling surface, and said composite material substrate wing assembly further comprising:
a composite material substrate wing assembly first surface;
a composite material substrate wing assembly second surface, said composite material substrate wing assembly first surface positioned proximate to the tooling surface;
a composite material substrate wing assembly first edge, said composite material substrate wing assembly first edge comprising at least one composite material substrate wing assembly first edge entry;
a composite material substrate wing assembly second edge, said composite material substrate wing assembly second edge comprising at least one composite material substrate wing assembly second edge entry;

orienting a separator layer onto composite material substrate wing assembly;
orienting a composite material top layer onto the separator layer;
at least partially curing the composite material substrate wing assembly and the composite material top layer;
separating the composite material substrate wing assembly and the composite material top layer;
orienting a composite material removal device proximate to the composite material substrate wing assembly second surface;
machining a plurality of recesses into the composite material substrate wing assembly;
removing a predetermined amount of material from the composite material substrate wing assembly to form the plurality of recesses in the composite material substrate wing assembly,
said plurality of recesses configured to extend from the composite material wing assembly second surface to a predetermined distance into the composite material substrate thickness to form a machined composite material substrate wing assembly, said predetermined distance equal to from about 50% to about 80% of the material substrate thickness;
said machined composite material substrate wing assembly comprising a machined composite material substrate wing assembly first surface and a machined composite material substrate wing assembly second surface; and

machining a plurality of channels into the composite material substrate wing assembly, with at least one of said plurality of channels extending from at least one of the recesses to at least one of: the material substrate first edge entry, the material substrate second edge entry, and another recess;
wherein the composite material substrate wing assembly is configured to form an internal reinforcing structural component, said internal reinforcing structural component further comprising a predetermined tailorable material stiffness, said predetermined tailorable material stiffness determined according to ASTM C297.

2. The method of claim 1 , wherein at least one of the plurality of channels extend from a recess to at least one of the composite material substrate wing assembly first edge entry or the composite material substrate second edge entry.

3. The method of claim 1 , wherein at least one of the plurality of recesses is in communication with at least one of: the composite material substrate wing assembly first edge entry; the composite material substrate wing assembly second edge entry; and another recess.

4. The method of claim 1 , wherein at least one recess comprises a geometric profile, said geometric profile comprising: at least one of: a rectangle; a circle; an ellipse; a triangle; a polygon; an irregular shape; or combinations thereof.

5. The method of claim 1 , wherein the composite material removal device is a laser.

6. The method of claim 1 , wherein the composite material removal device is a CNC machine.

7. The method of claim 1 , wherein the composite material removal device is in communication with a controller.

8. The method of claim 1 , further comprising:
machining the plurality of recesses into the composite material substrate wing assembly in a predetermined pattern.

9. The method of claim 1 , further comprising:
machining at least one of the plurality of recesses into an at least partially cured composite material substrate wing assembly, said at least one of the plurality of recesses configured to extend from the composite material substrate wing assembly second surface to a predetermined distance into the composite material substrate thickness to form the machined composite material substrate wing assembly, said machined composite material substrate wing assembly comprising a machined composite material substrate wing assembly first surface and a machined composite material substrate wing assembly second surface;
applying an adhesive material to at least one of the composite material top layer or the machined composite material substrate wing assembly second surface;
orienting the composite material top layer onto the machined composite material substrate wing assembly second surface; and
bonding the composite material top layer to the machined composite material substrate wing assembly second surface.

10. The method of claim 9 , further comprising;
applying a negative pressure to the at least one of the plurality of recesses; and
curing the composite material substrate wing assembly.

11. The method of claim 10 , further comprising:
sealing at least one of the composite material substrate wing assembly first edge entry and the composite material substrate wing assembly second edge entry.

12. A wing assembly comprising: a wing panel, said wing panel comprising:
a material substrate comprising a material substrate thickness, said material substrate further comprising:
a material substrate first surface;
a material substrate second surface, said material substrate second surface comprising a plurality of recesses, said recesses configured to extend from the material second surface to a predetermined distance into the material substrate thickness from the material substrate second surface, said predetermined distance equal to from about 50% to about 80% of the material substrate thickness;
a material substrate first edge, said material substrate first edge comprising at least one material substrate first edge entry;
a material substrate second edge, said material substrate second edge comprising at least one material substrate second edge entry;
a plurality of channels in the material substrate, with at least one of said plurality of channels extending from at least one of the recesses to at least one of: the material substrate first edge entry, the material substrate second edge entry, and another recess; and
a top layer material bonded to the material substrate second surface to form a wing assembly, and to further form an internal pathway in the material substrate, said internal pathway bounded by the material substrate second surface and the top layer material;

wherein at least one of the material substrate and the top layer material comprises a composite material; and
wherein the material substrate further comprises a predetermined tailorable material stiffness, said predetermined tailorable material stiffness determined according to ASTM C297.

13. The wing assembly of claim 12 , wherein the material substrate comprises a material substrate interior pressure, said material substrate exposed to a material substrate exterior pressure, and wherein the material substrate interior pressure is substantially equivalent to the material substrate exterior pressure, said material substrate exterior pressure substantially equivalent to a

Assignees

• Boeing Co

Inventors

• Anderson David M
• Baucum Jeffrey P

Classifications

• B32B2260/021

  Fibrous or filamentary layer · CPC title

• B64F5/10 Primary

  Manufacturing or assembling aircraft, e.g. jigs therefor · CPC title

• B32B15/08

  of synthetic resin · CPC title

• B32B5/26

  another layer {next to it} also being fibrous or filamentary {(relative arrangement of fibres or filaments of different layers B32B5/12)} · CPC title

• B32B2262/106

  Carbon fibres, e.g. graphite fibres · CPC title