Optimized Cross-Ply Orientation in Composite Laminates

• US 20130330503A1
• Filed: 06/08/2012
• Published: 12/12/2013
• Est. Priority Date: 06/08/2012
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Kismarton, Max U.
• Granted Patent: US 9,289,949 B2
• Time in Patent Office: Days
• US Class Current: 428/113
• 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 having a primary axis of loading, comprising:

a plurality of resin plies each reinforced with unidirectional fibers, and including cross-plies have fiber orientations optimized to resist bending and torsional loads along the primary axis of loading.

2. The composite laminate of claim 1, wherein the plurality of resin plies includes:

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

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

the cross-plies each having ±

θ

degree fiber orientations relative to the primary axis of loading, where θ

is within a range of approximately 10 degrees and 43 degrees.

3. The composite laminate of claim 2, wherein θ

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

4. The composite laminate of claim 2, wherein θ

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

5. The composite laminate of claim 2, wherein the θ

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

6. The composite laminate of claim 2, wherein the θ

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

7. The composite laminate of claim 2, wherein the θ

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

8. A composite laminate aircraft skin having a primary axis of loading, comprising:

a first group of fiber reinforced resin plies having a fiber orientation substantially parallel to the primary axis of loading;

a second group of fiber reinforced resin plies having a fiber orientation substantially perpendicular to the primary axis of loading; and

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

θ

degree fiber orientation relative to the primary axis of loading, where θ

is optimized to resist bending loads and torsional loads applied to the skin.

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

is within a range of approximately 33 and 43 degrees.

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

is within a range of approximately 35 and 40 degrees.

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

varies across the skin along the primary axis of loading.

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

the 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. A composite laminate aircraft skin having a primary axis of loading, comprising:

at least one group of fiber reinforced resin cross-plies having fiber orientations that vary along the primary axis of loading.

14. The composite laminate aircraft skin of claim 13, wherein the fiber orientations of the cross-plies vary within a range of approximately 10 and 43 degrees.

15. The composite laminate aircraft skin of claim 13, further comprising:

plies having fiber orientations of approximately 90 degrees relative to the primary axis of loading, and

plies having fiber orientations of approximately 0 degrees relative to the primary axis of loading.

16. A method of laying up a composite aircraft skin having a primary axis of loading, comprising:

laying up cross-plies of resin reinforced with unidirectional fibers, including orienting the cross-plies at angles that vary along the primary axis of loading.

17. The method of claim 16, wherein the angles are within the range of approximately +10 and +43 degrees, and −

10 and −

43 degrees relative to the primary axis of loading.

18. The method of claim 16, wherein orienting the cross-plies includes selecting orientation angles for the cross plies that are substantially matched to loads imposed on the wing skin at multiple locations along the primary axis of loading.

19. A method of fabricating a composite aircraft wing skin having a primary axis of loading, comprising:

laying up a plurality of resin plies each reinforced with unidirectional fibers having a fiber orientation substantially parallel to the primary axis of loading;

laying up a plurality of resin plies each reinforced with unidirectional fibers having a fiber orientation substantially orthogonal to the primary axis of loading; and

laying up a plurality of resin cross-plies each reinforced with unidirectional fibers and having an angular fiber orientation ±

θ

, including optimizing the angular fiber orientation ±

θ

to substantially match the loading on the wing skin.

20. The method of claim 19, wherein optimizing the angular orientation ±

θ

is performed at each of a plurality of locations along the primary axis of loading.

21. The method of claim 19, wherein optimizing the angular orientation ±

θ

includes selecting a fiber orientation angle within the range of approximately 10 and 43 degrees.

22. A method of reducing the weight of a composite wing skin formed of fiber reinforced plies that include 0 degree and 90 degree plies traversed by cross-plies, comprising:

reducing the number of 0 degree plies required to resist loads on the wing skin by optimizing the angular orientation of the cross plies.

23. The method of claim 22, wherein optimizing the angular orientation of the cross-plies includes selecting an angular orientation ±

θ

for the cross-plies, where ±

θ

is within the range of approximately 10 and 43 degrees.

24. The method of claim 22, wherein the 90 degree plies extend substantially parallel to a primary axis of loading of the wing skin, and optimizing the angular orientation of the cross plies is performed at more than one location in the span-wise direction of the wing skin along the primary axis of loading.