Optimized cross-ply orientation in composite laminates

• US 10,000,025 B2
• Filed: 03/21/2016
• Issued: 06/19/2018
• 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:
  US15/076,405
• Publication Number:
  US 20160200054A1
• Time in Patent Office: 820 Days
• Field of Search: 428113
• 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 method of laying up plies in an aircraft wing skin having an axis of loading in plane with the aircraft wing skin, comprising:

laying up 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 from a root to a tip of the aircraft wing skin;

laying up a first set of cross-plies having a second fiber orientation of approximately 43 degrees relative to the axis of loading and the first fiber orientation, the first set of cross-plies extending from the root to a first location on the aircraft wing skin;

laying up a second set of cross plies having a third fiber orientation relative to the axis of loading in a first range of approximately 40-43 degrees and extending from the first location to a second location approximately midway between the first location and the tip of the aircraft wing skin; and

laying up a third set of cross plies having a fourth fiber orientation relative to the axis of loading in a second range that decreases from approximately 40 degrees to approximately 25 degrees from the second location to the tip of the aircraft wing skin;

wherein the plies are resin and are reinforced with unidirectional fibers.

2. The method of claim 1, wherein the second fiber orientation of the first set of cross-plies is selected based on a loading of the aircraft wing skin along the axis of loading.

3. The method of claim 1, wherein laying up the first set of cross-plies includes selecting orientation angles for the first set of cross-plies that are based on loads imposed on the aircraft wing skin at multiple locations along the axis of loading.

4. A method of fabricating a composite aircraft wing skin having an axis of loading in plane with the composite aircraft wing skin, the method comprising:

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

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

laying up a third plurality of resin cross-plies each reinforced with unidirectional fibers and having an angular fiber orientation relative to the axis of loading, including optimizing the angular fiber orientation based on a loading on the composite aircraft wing skin;

wherein optimizing the angular fiber orientation includes selecting a fiber orientation angle relative to the axis of loading that decreases from approximately 40 degrees to approximately 25 degrees from a selected location on the composite aircraft wing skin to a tip of the composite aircraft wing skin.

5. The method of claim 4, wherein optimizing the angular fiber orientation is performed at each of a plurality of locations extending from a root of the composite aircraft wing skin to the tip of the composite aircraft wing skin.

6. The method of claim 4, wherein a number of angular fiber orientations of the first plurality of resin plies, the second plurality of resin plies, and the third plurality of resin cross-plies vary from a root to the tip of the composite aircraft wing skin.

7. The method of claim 4, wherein the angular fiber orientation of the cross-plies varies within an area of the composite aircraft wing skin.

8. The method of claim 4, wherein the angular fiber orientation of the cross-plies is selected based on a loading of the composite aircraft wing skin from the root to the tip of the composite aircraft wing skin.

9. The method of claim 4, wherein optimizing the angular fiber orientation based on the loading on the composite aircraft wing skin includes selecting orientation angles for the cross-plies based on loads imposed on the composite aircraft wing skin at multiple locations along the axis of loading.