Composite structure splice and method

Patent metadata

Field	Value
Publication number	US-11845236-B2
Application number	US-201815918508-A
Country	US
Kind code	B2
Filing date	Mar 12, 2018
Priority date	Mar 12, 2018
Publication date	Dec 19, 2023
Grant date	Dec 19, 2023

Abstract

A method for fabricating a composite structure. A first section for the composite structure is formed in which the first section has a first end with a chevron shape, wherein first composite layers in the first section has a first step pattern at the first end. A second section for the composite structure is formed in which the second section has a second end with a counterpart shape to the chevron shape and in which second composite layers in the second section have a second step pattern at the second end. The first end the second end are positioned such that a first composite layer in the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at a splice location.

First claim

What is claimed is:

1. A method of forming a step lap splice for a composite structure, the method comprising:
forming a chevron shape in each layer of a first section to create a first step lap pattern in the first section in which a plurality of chevrons is formed in the first section;
forming a counterpart shape in each layer of a second section to create a second step lap pattern in the second section in which a plurality of counterparts is formed in the second section; and
overlapping each layer with the chevron shape with a corresponding counterpart shape in the second section to form the step lap splice.

2. The method of claim 1 , wherein locations of overlaps and gaps in the step lap splice in the composite structure reduces undesired wrinkling in the composite structure.

3. The method of claim 1 , wherein the step lap splice is placed in a number of locations in the composite structure, wherein the number of locations is selected to reduce undesired wrinkling in the composite structure.

4. The method of claim 1 , wherein robotic arms are utilized to steer at least one of the first section and the second section to overlap each layer with the chevron shape with the corresponding counterpart shape in the second section to form the step lap splice.

5. The method of claim 1 , wherein the plurality of chevrons and the plurality of counterparts mitigate crack propagation in at least one of the composite structure or the step lap splice.

6. The method of claim 1 further comprising:
selecting a splice location for the step lap splice in the composite structure to reduce buckling inducing compressive stress in the composite structure.

7. A portion of an aircraft assembled according to the method of claim 1 .

8. A method for mitigating crack propagation in a composite structure, the method comprising:
identifying a splice location for the composite structure to reduce stress in the composite structure; and
forming a step lap splice at the splice location in which the step lap splice has a chevron shape in each layer of a first section that forms a first step lap pattern in the first section in which a plurality of chevrons is in the first section; in which the step lap splice has a counterpart shape in each layer of a second section that forms a second step lap pattern in the second section in which a plurality of counterparts is in the second section; and in which each layer with the chevron shape overlaps a corresponding counterpart shape for another layer in the second section to form the step lap splice.

9. The method of claim 8 , wherein forming the step lap splice at the splice location reduces at least one of undesired wrinkling or stress in the composite structure at the splice location.

10. A portion of an aircraft assembled according to the method of claim 8 .

11. A method for fabricating a composite structure, the method comprising:
forming a first section for the composite structure in which the first section has a first end with a chevron shape, wherein first composite layers in the first section has a first step pattern at the first end;
forming a second section for the composite structure in which the second section has a second end with a counterpart shape to the chevron shape and in which second composite layers in the second section have a second step pattern at the second end; and
positioning the first end and the second end such that the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at a splice location.

12. The method of claim 11 , wherein undesired wrinkling for the composite structure is reduced at and around the splice location.

13. The method of claim 11 , wherein the composite structure is a stringer, and wherein the first section and the second section are substantially planar and further comprising:
forming the first section and the second section into a cross-sectional shape created by a tool having an upper die and lower die.

14. The method of claim 13 , wherein the cross-sectional shape is a hat shape, a U-shape, or a C-shape.

15. The method of claim 11 further comprising:
placing the first section and the second section on a tool for the composite structure; and
curing the first section and the second section to form the composite structure.

16. The method of claim 11 , wherein the chevron shape comprises a group of chevrons.

17. The method of claim 11 , wherein positioning the first end and the second end such that the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at the splice location comprises:
positioning the first end and the second end such that the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at the splice location, wherein at least one of a gap or an overlap is present between the first composite layers and the second composite layers such that stress is reduced at the splice location.

18. The method of claim 11 , wherein the composite structure is a stringer and further comprises:
attaching the stringer to a composite barrel section of a fuselage for an aircraft.

19. The method of claim 11 , wherein positioning the first end and the second end such that the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at the splice location comprises:
positioning the first end and the second end such that the first composite layers in the first step pattern overlap the second composite layers in the second step pattern to form a wedge splice at the splice location.

20. The method of claim 11 , wherein the first composite layers are a first prepreg structure and the second composite layers are a second prepreg structure.

21. The method of claim 11 , wherein the composite structure is selected from the group consisting of a stiffener, a stringer, a longeron, and a beam.

22. A portion of an aircraft assembled according to the method of claim 11 .

23. A method for fabricating a composite structure, the method comprising:
forming a first section for the composite structure in which the first section has a first end with a chevron shape in which first composite layers in the first section have a first step pattern at the first end, and in which the chevron shape comprises a group of chevrons;
forming a second section for the composite structure in which the second section has a second end with a counterpart shape to the chevron shape, and in which second composite layers in the second section have a second step pattern at the second end in which the first section and the second section are substantially planar;
forming the first section and the second section into a cross-sectional shape created by a tool having an upper die and lower die;
positioning the first end and the second end on a tool such that the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at a splice location to form a lap splice joint in which at least one of a gap or an overlap is present between the first composite layers and the second composite layers such that stress is reduced at the lap splice joint and undesired wrinkling for the composite structure is reduced at the splice location; and
curing the first section and the second section to form the composite structure.

24. A portion of an aircraft as

Assignees

• Boeing Co

Inventors

• Kendall James R
• Suriyaarachchi Raviendra Sidath
• Shaw Paul D

Classifications

• B29D99/0003 Primary

  Producing profiled members, e.g. beams · CPC title

• B29C70/304

  In-plane lamination by juxtaposing or interleaving of plies, e.g. scarf joining · CPC title

• B29C70/34

  and shaping or impregnating by compression {, i.e. combined with compressing after the lay-up operation} · CPC title

• B29K2105/089

  fabric · CPC title

• B29K2105/0881

  unidirectional · CPC title