Spliced laminate for aircraft fuselage

• US 5,429,326 A
• Filed: 07/09/1992
• Issued: 07/04/1995
• Est. Priority Date: 07/09/1992
• Status: Expired due to Term

• Current Assignee: Alcoa Incorporated (Howmet Aerospace, Inc.)
• Original Assignee: Structural Laminates Company (Howmet Aerospace, Inc.)
• Inventors: Garesche, Carl E., Gunnink, Jan W., Roebroeks, Gerandus H. J. J., Greidanus, Buwe V. W., Oost, Rob C. V.
• Primary Examiner(s): Hill, Mitchell J.
• Assistant Examiner(s): ELLIS, CHRISTOPHER P
• Application Number:
  US07/910,938
• Time in Patent Office: 1,090 Days
• Field of Search: 244/119, 244/158, 244/120, 244/123, 244/124, 244/131, 244/132, 244/133, 244/118.2, 428/189, 428/77, 428/213, 428/215, 428/216, 156/304.1, 156/157, 156/547, 156/549, 156/502, 280/610
• US Class Current: 244/133
• CPC Class Codes: B32B 15/08 of synthetic resin
• CPC Class Codes: B32B 15/18 comprising iron or steel B32B15/011, B32B15/012 and B32B15/013 take precedence
• CPC Class Codes: B32B 15/20 comprising aluminium or copper B32B15/016 and B32B15/017 take precedence
• CPC Class Codes: B32B 2309/105 Thickness
• CPC Class Codes: B32B 2605/18 Aircraft
• CPC Class Codes: B32B 3/10 characterised by a discontinuous layer, i.e. formed of separate pieces of material
• CPC Class Codes: B32B 7/12 using interposed adhesives or interposed materials with bonding properties
• CPC Class Codes: B64C 1/12 Construction or attachment of skin panels
• CPC Class Codes: Y10T 156/1722 Means applying fluent adhesive or adhesive activator material between layers
• CPC Class Codes: Y10T 428/24752 Laterally noncoextensive components

Abstract

A laminated body panel for aircraft applications comprises first and second metal layers and an adhesive layer. First and second sections in each metal layer are generally coplanar and separated by a splice line. A first splice line in a first metal layer is generally parallel to a second splice line in a second metal layer, and spaced laterally from the second splice line. An adhesive layer between the metal layers preferably contains reinforcing fibers bridging across the splice lines. An aircraft fuselage or wing or empennage made with spliced body panels of the invention has fewer joints and weighs less than one made with narrower body panels containing unspliced metal layers.

Claims

1. A laminated panel comprising at least:

(a) a first metal layer having a thickness of less than about 1 mm and comprising a first section, a second section generally coplanar with the first section and a first splice line between the first and second sections;

(b) a second metal layer having a thickness of less than about 1 mm and comprising a first section, a second section generally coplanar with the first section and a second splice line between the first and second sections, the second splice line being spaced laterally from the first splice line by at least several centimeters; and

(c) a fiber reinforced adhesive layer between the first and second metal layers.

2. The panel in accordance with claim 1 comprising 2 to about 20 metal layers and 1 to about 19 adhesive layers.

3. The panel in accordance with claim 1 wherein said first and second splice lines are generally parallel.

4. The panel in accordance with claim 1 wherein said metal layers have first and second lateral edges on opposite lateral sides of said splice lines, said first and second lateral edges being separated by a width of at least about 165 cm.

5. The panel in accordance with claim 4 wherein said width amounts to at least about 200 cm.

6. The panel in accordance with claim 4 wherein said width is at most 800 cm.

7. The panel in accordance with claim 4 wherein said width is between 250 and 400 cm.

8. The panel in accordance with claim 1 having a rectangular shape with a length and a width and wherein the splice lines are generally parallel to the longer of the length and the width.

9. The panel in accordance with claim 1 wherein at least a substantial portion of said fibers extends generally perpendicular to said first and second splice lines.

10. The panel in accordance with claim 9 wherein said portion of the fibers bridge across said first and second splice lines.

11. The panel in accordance with claim 1 wherein at least a substantial portion of said fibers extends generally parallel to said first and second splice lines.

12. The panel in accordance with claim 1 wherein a first group of said fibers extends generally parallel to said first and second splice lines and a second group of said fibers extends generally perpendicular to said first and second splice lines.

13. The panel in accordance with claim 1 wherein said fibers are generally continuous.

14. The panel in accordance with claim 1 wherein said first splice line extends generally perpendicular to said second splice line.

15. The panel in accordance with claim 1 wherein the metal layers each have a thickness of 0.1-0.9 mm.

16. The panel in accordance with claim 1 wherein the metal layers each have a thickness of 0.2-0.5 mm.

17. The panel in accordance with claim 1 wherein the thickness of each fiber reinforced adhesive layer is generally of the same order of magnitude as the thickness of individual metal layers.

18. The panel in accordance with claim 1 wherein the fibers constitute 35-75 vol. % of the total volume of adhesive and fibers combined.

19. The panel in accordance with claim 1 wherein said metal layers are made from one or more metals selected from the group consisting of aluminum alloys, steel alloys, copper alloys, zinc alloys, titanium alloys, magnesium alloys and aluminum alloy matrix composites.

20. The panel in accordance with claim 1 wherein said metal layers are made from a metal having a tensile strength of more than 0.20 GPa.

21. The panel in accordance with claim 1 wherein said fibers are made from one or more materials selected from the group consisting of glass, E-glass, R-glass, S2-glass, aramids, poly-paraphenyleneterephthalamide and carbon.

22. The panel in accordance with claim 1 wherein said fibers are made from a material having a modulus of elasticity of 60-650 GPa and an elongation at break of 0.2-8%.

23. The panel in accordance with claim 13 wherein said fibers are generally continuous filaments each having a diameter of 3-30 μ

m.

24. The panel in accordance with claim 1 wherein one or more metal layers each comprise 3-15 coplanar sections with parallel splice lines between the sections.

25. An aircraft fuselage or empennage which is at least partly built up of the laminated panel according to claim 1.

26. An aircraft wing which is at least partly built up of the laminated panel according to claim 1.

27. A laminated panel comprising at least:

(a) a first aluminum alloy layer having a thickness of less than about 1 mm and comprising a first section, a second section generally coplanar with the first section and a first splice line between the first and second sections;

(b) a second aluminum alloy layer having a thickness of less than about 1 mm and comprising a first section, a second section generally coplanar with the first section and a second splice line between the first and second sections, the second splice line being generally parallel to the first splice line and spaced laterally therefrom by at least several centimeters; and

(c) an adhesive layer between the first and second aluminum alloy layers, said adhesive layer comprising a synthetic polymer reinforced with fibers comprising a material selected from the group consisting of glass, E-glass, R-glass, S2-glass, aramids, poly-paraphenyleneterephthalamide and carbon.

28. The panel in accordance with claim 27 wherein said first and second splice lines are generally parallel.

29. The panel in accordance with claim 27 wherein said metal layers have first and second lateral edges on opposite lateral sides of said splice lines, said first and second lateral edges being separated by a width of at least about 165 cm.

30. The panel in accordance with claim 29 wherein said width amounts to at least about 200 cm.

31. The panel in accordance with claim 29 wherein said width is at most 800 cm.

32. The panel in accordance with claim 29 wherein said width is between 250 and 400 cm.

33. The panel in accordance with claim 27 having a rectangular shape with a width and a length longer than the width, said splice lines being generally parallel to the length.

34. The panel in accordance with claim 27 wherein at least a substantial portion of said fibers extends generally perpendicular to said first and second splice lines.

35. The panel in accordance with claim 34 wherein said substantial portion of the fibers bridge across said first and second splice lines.

36. An aircraft fuselage or empennage at least partially built up of the laminated panel of claim 27.

37. An aircraft wing at least partially built up of the laminated panel of claim 27.