Optimized cross-ply orientation in composite laminates

• US 9,289,949 B2
• Filed: 06/08/2012
• Issued: 03/22/2016
• Est. Priority Date: 06/08/2012
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Kismarton, Max U.
• Primary Examiner(s): O'HERN, BRENT T
• Application Number:
  US13/491,784
• Publication Number:
  US 20130330503A1
• Time in Patent Office: 1,383 Days
• Field of Search: 428/113, 428/98, 428/105
• US Class Current: 1/1
• CPC Class Codes: B29C 70/202 arranged in parallel planes or structures of fibres crossing at substantial angles, e.g. cross-moulding compound [XMC] B29C70/207 takes precedence
• CPC Class Codes: B29C 70/30 Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
• CPC Class Codes: B29K 2105/0881 unidirectional
• CPC Class Codes: B29L 2031/3076 Aircrafts blades, propellers B29L2031/08
• CPC Class Codes: B29L 2031/3085 Wings
• CPC Class Codes: B32B 2307/544 Torsion strength; Torsion stiffness
• CPC Class Codes: B32B 2307/546 Flexural strength; Flexion stiffness
• CPC Class Codes: B32B 2605/18 Aircraft blades, propellers B32B2603/00
• CPC Class Codes: B32B 5/12 characterised by the relative arrangement of fibres or filaments of different layers, e.g. the fibres or filaments being parallel or perpendicular to each other
• CPC Class Codes: B32B 5/26 another layer next to it also being fibrous or filamentary relative arrangement of fibres or filaments of different layers B32B5/12
• CPC Class Codes: Y02T 50/40 Weight reduction
• CPC Class Codes: Y10T 156/10 Methods of surface bonding and/or assembly therefor
• CPC Class Codes: Y10T 428/24124 Fibers

Abstract

A composite laminate has a primary axis of loading and comprises a plurality resin plies each reinforced with unidirectional fibers. The laminate includes cross-plies with fiber orientations optimized to resist bending and torsional loads along the primary axis of loading.

Claims

1. A composite laminate aircraft wing skin having an axis of loading in-plane with the aircraft wing skin, the composite laminate aircraft skin comprising:

a plurality of resin plies in the aircraft wing skin reinforced with unidirectional fibers, wherein the plurality of resin plies comprises;

a first set of plies having a substantially straight first fiber orientation of substantially 0 degrees relative to the axis of loading, the first set of plies extending a length of the wing, wherein the axis of loading substantially bisects the wing; and

at least one first set of cross-plies having a second fiber orientation of ±

θ

degrees relative to the axis of loading and the first fiber orientation, wherein 0 is determined while the composite laminate aircraft wing skin is in a static position, where θ

is greater than or equal to approximately 25 degrees and less than or equal to approximately 43 degrees.

2. The composite laminate aircraft wing skin of claim 1, wherein θ

is within a range of approximately 33 degrees and 43 degrees.

3. The composite laminate aircraft wing skin of claim 1, wherein θ

is within a range of approximately 35 degrees and 40 degrees.

4. The composite laminate aircraft wing skin of claim 1, wherein the θ

degree fiber orientations of the cross-plies vary along the axis of loading.

5. The composite laminate aircraft wing skin of claim 1, wherein the θ

degree fiber orientations of the cross-plies vary within an area of the composite laminate aircraft wing skin.

6. The composite laminate aircraft wing skin of claim 1, wherein the θ

degree fiber orientations of the cross-plies are selected to match loading of composite laminate aircraft wing skin along the axis of loading.

7. The composite laminate aircraft wing skin of claim 1, wherein the plurality of resin plies further includes:

at least one resin ply reinforced with unidirectional fibers having a substantially 0 degree fiber orientation relative to the axis of loading, and

at least one resin ply reinforced with unidirectional fibers having a substantially 90 degree fiber orientation relative to the axis of loading.

8. A composite laminate aircraft wing skin having an axis of loading in-plane with the aircraft wing skin, the composite laminate aircraft skin comprising:

a first group of fiber reinforced resin plies extending a length of the wing and having a first fiber orientation substantially straight and parallel to the axis of loading of the composite laminate aircraft skin, wherein the axis of loading substantially bisects the wing, wherein the first fiber orientation is determined while the composite laminate aircraft wing skin is in a static position;

a second group of fiber reinforced resin plies having second a fiber orientation substantially perpendicular to the axis of loading of the composite laminate aircraft skin, wherein the second fiber orientation is determined while the composite laminate aircraft wing skin is in a static position; and

a third group of fiber reinforced resin cross-plies extending transverse to the plies in the first and second groups thereof and having a ±

θ

degree fiber orientation relative to the axis of loading of the composite laminate aircraft skin and the first fiber orientation, wherein θ

is determined while the composite laminate aircraft wing skin is in a static position, where θ

is optimized to resist bending loads and torsional loads applied to the skin, wherein θ

is greater than or equal to approximately 25 degrees and less than approximately 43 degrees.

9. The composite laminate aircraft wing skin of claim 8, wherein θ

is within a range of approximately 33 and 43 degrees.

10. The composite laminate aircraft wing skin of claim 8, wherein θ

is within a range of approximately 35 and 40 degrees.

11. The composite laminate aircraft wing skin of claim 8, wherein θ

varies across the composite laminate aircraft skin along the axis of loading.

12. The composite laminate aircraft wing skin of claim 8 wherein:

the composite laminate aircraft skin is a wing skin having a wing root and a wing tip,

θ

is approximately 43 degrees in the area of the wing root, and

θ

is approximately 10 degrees at the wing tip.

13. The composite laminate aircraft wing skin of claim 8, wherein the first group, the second group, and the third group of fiber reinforced resin plies comprise a total number of plies, and wherein the second group comprises approximately 10% of the total number of plies.

14. A composite laminate aircraft wing skin having an axis of loading in-plane with the composite laminate aircraft wing skin, comprising:

a first number of fiber reinforced resin plies extending a length of the wing having substantially straight first fiber orientations of approximately 0 degrees relative to the axis of loading of the composite laminate aircraft skin, wherein the axis of loading substantially bisects the wing, wherein the fiber orientations are determined while the composite laminate aircraft wing skin is in a static position;

a second number of fiber reinforced resin plies having second fiber orientations of approximately 90 degrees relative to the axis of loading of the composite laminate aircraft skin, wherein the fiber orientations are determined while the composite laminate aircraft wing skin is in a static position, and

a third number of fiber reinforced resin cross-plies having ±

θ

degree fiber orientations that vary relative to the axis of loading of the composite laminate aircraft skin and the first fiber orientations, where θ

is greater than or equal to approximately 10 degrees and less than or equal to approximately 43 degrees, wherein θ

is determined while the composite laminate aircraft wing skin is in a static position.

15. The composite laminate aircraft wing skin of claim 14, wherein θ

varies within a range of approximately 10 and 43 degrees along the axis of loading.

16. The composite laminate aircraft wing skin of claim 14, wherein θ

decreases in a step-wise manner along the axis of loading.

17. A composite laminate aircraft wing skin having a primary axis of loading in-plane with the composite laminate aircraft wing skin, comprising:

a plurality of resin plies each reinforced with unidirectional fibers, wherein the plurality of resin plies comprises;

a first set of plies having a substantially straight first fiber orientation of substantially 0 degrees relative to the axis of loading, the first set of plies extending a length of the wing,

wherein the axis of loading substantially bisects the wing; and

at least one first set of cross-plies having a second fiber orientation of ±

θ

degrees relative to the primary axis of loading and the first fiber orientation, where θ

is greater than or equal to approximately 25 degrees and less than or equal to approximately 43 degrees, wherein θ

is determined while the composite laminate aircraft wing skin is in a static position and optimized to resist bending and torsional loads along the primary axis of loading of the composite laminate aircraft wing skin, wherein the optimized fiber orientations vary in a span-wise direction along the aircraft wing skin.

18. The composite laminate aircraft wing skin of claim 17, wherein θ

decreases linearly in a span-wise direction along the aircraft skin.

19. The composite laminate aircraft wing skin of claim 17, wherein θ

decreases non-linearly in a span-wise direction along the aircraft skin.

20. The composite laminate aircraft wing skin of claim 17, wherein θ

varies within a range of approximately 10 to 43 degrees from a wing tip to a wing root.