Ply blending and stacking sequence

• US 10,195,812 B2
• Filed: 08/26/2015
• Issued: 02/05/2019
• Est. Priority Date: 08/26/2015
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Blom, Adriana Willempje, Kang, Laura Sumi, Epton, Michael Andrew
• Primary Examiner(s): O'Hern, Brent T
• Application Number:
  US14/836,369
• Publication Number:
  US 20170057195A1
• Time in Patent Office: 1,259 Days
• Field of Search: 428113
• CPC Class Codes: B32B 2307/50 having particular mechanical properties
• CPC Class Codes: B32B 2307/732 Dimensional properties
• CPC Class Codes: B32B 2605/18 Aircraft blades, propellers B32B2603/00
• CPC Class Codes: B32B 2607/00 Walls, panels
• CPC Class Codes: B32B 3/10 characterised by a discontinuous layer, i.e. formed of separate pieces of material
• CPC Class Codes: B32B 5/02 characterised by structural features of a fibrous or filamentary layer layer formed of metallic wires B32B15/02; layer formed of natural mineral fibres B32B19/02; layer formed of wood fibres B32B21/02
• CPC Class Codes: B64C 3/26 Construction, shape, or attachment of separate skins, e.g. panels
• CPC Class Codes: G06F 2113/24 Sheet material
• CPC Class Codes: G06F 2113/26 Composites
• CPC Class Codes: G06F 30/15 Vehicle, aircraft or watercraft design
• CPC Class Codes: G06F 30/20 Design optimisation, verification or simulation optimisation, verification or simulation of circuit designs G06F30/30

Abstract

A composite laminate is presented. The composite laminate comprises a plurality of plies having non-traditional angles. Each of the plurality of plies has a respective shape configured to minimize a total length of the ply boundary over all of the plurality of plies. The composite laminate has a stiffness equal to a desired stiffness of a traditional composite laminate.

Claims

1. A composite laminate part having a designated reference axis that extends along the composite laminate part in a same plane or a parallel plane as a longitudinal direction of fibers within a ply in a corresponding plane, the composite laminate part comprising:

a plurality of plies that together have a specified thickness and stiffness distribution by having non-traditional fiber angles for fibers of the plurality of plies, wherein the non-traditional fiber angles consist of fiber angles other than zero degrees, ninety degrees, or plus or minus forty-five degrees relative to the reference axis, wherein corresponding fiber angles of any two adjacent plies are different relative to each other, wherein each of the plurality of plies has a respective shape configured to minimize a total ply boundary length over all of the plurality of plies, wherein a ply boundary is a boundary between two adjacent panels in one ply sequence in which a given ply is present in one panel and the given ply is absent in another panel in the same ply sequence, wherein a ply boundary length is a first sum of lengths of all boundaries of the given ply, and wherein the total ply boundary length is a second sum of all ply boundary lengths of the plurality of plies, wherein the respective shape was determined for each ply using an objective function by grouping panels of the plurality of plies with a same quantity of plies and a same set of lamination parameters to form a plurality of groups of panels and to create ply sequences for the plurality of groups of panels, wherein the plurality of groups of panels have the ply sequences, wherein at least some of the plurality of groups of panels are different in shape relative to others of the plurality of groups of panels, wherein the composite laminate part has a stiffness equal to a desired stiffness of a traditional composite laminate part, wherein the desired stiffness comprises lamination parameters above a particular threshold expected for a different plurality of plies having other fibers oriented at traditional angles relative to the reference axis, wherein traditional angles consist of fiber angles of only zero degrees, ninety degrees, or plus or minus forty-five degrees for the different plurality of plies relative to the reference axis.

2. The composite laminate part of claim 1, wherein the plurality of plies has rectangular edges.

3. The composite laminate part of claim 1, wherein the composite laminate part has a thickness less than a thickness of the traditional composite laminate part.

4. The composite laminate part of claim 1, wherein each group of the plurality of groups contains a respective portion of a first surface of the composite laminate part, and wherein a respective layup of the each group of the plurality of groups is different than a respective layup of each other group of the plurality of groups.

5. The composite laminate part of claim 4, wherein a quantity of the ply sequences is equal to at least a quantity of the plurality of groups times a quantity of lamination parameters.

6. The composite laminate part of claim 5, wherein for any two groups in the plurality of groups having a thickness difference of fewer plies than the quantity of lamination parameters, the two groups have a quantity of plies that are present in only one of the two groups equal to at least the quantity of lamination parameters.

7. A composite laminate part having a designated reference axis that extends along the composite laminate part in a same plane or a parallel plane as a longitudinal direction of fibers within a ply in a corresponding plane, the composite laminate part comprising:

a plurality of panels formed of a plurality of plies having non-traditional fiber angles, wherein the non-traditional fiber angles consist of fiber angles other than zero degrees, ninety degrees, or plus or minus forty-five degrees, relative to the reference axis, wherein corresponding fiber angles relative to any two plies in the plurality of plies are different relative to each other, and wherein each panel shares at least one ply with an adjacent panel.

8. The composite laminate part of claim 7, wherein each of the plurality of plies has a respective shape configured to optimize blending across the plurality of panels.

9. The composite laminate part of claim 7, wherein each of the plurality of plies has a respective shape configured to minimize the total ply boundary length over all of the plurality of panels, wherein a ply boundary is a boundary between two adjacent panels in one ply sequence in which a given ply is present in one panel and the given ply is absent in another panel in the same ply sequence, wherein a ply boundary length is a first sum of lengths of all boundaries of the given ply, and wherein the total ply boundary length is a second sum of all ply boundary lengths of the plurality of plies.

10. The composite laminate part of claim 7, wherein each ply of a first half of a panel of the plurality of panels has a different fiber angle value than each other ply of the first half of the panel.

11. The composite laminate part of claim 7, wherein a first panel of the plurality of panels is symmetric about a mid-plane of the composite laminate part.

12. The composite laminate part of claim 7, wherein the composite laminate part has a thickness less than a thickness of a traditional composite laminate part having a same stiffness as the composite laminate part, wherein the traditional composite laminate part has fiber angles of only zero degrees, ninety degrees, or plus or minus forty-five degrees relative to the reference axis.

13. The composite laminate part of claim 7, wherein the composite laminate part is a wing skin for an aircraft wing, and wherein a width of a panel of the plurality of panels is a distance between two adjacent stringers of the aircraft wing.

14. The composite laminate part of claim 7, wherein at least one panel of the plurality of panels is rectangular.