Titanium-polymer hybrid laminates

• US 6,114,050 A
• Filed: 12/29/1998
• Issued: 09/05/2000
• Est. Priority Date: 01/11/1996
• Status: Expired due to Term

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Hoffman, Daniel J., Evans, David W., Westre, Willard N., Rothschilds, Robert J., Rogalski, Mark E., Allen-Lilly, Heather C., Ayers, Donald J., Cregger, Samuel E., Grande, Donald L.
• Primary Examiner(s): Jones, Deborah
• Assistant Examiner(s): LA VILLA, MICHAEL EUGENE
• Application Number:
  US09/222,371
• Time in Patent Office: 616 Days
• Field of Search: 428/596, 428/606, 428/608, 428/615, 428/292, 428/293, 428/458, 428/660, 428/220, 428/332, 428/457, 428/626, 428/118, 428/77, 428/189, 428/334, 428/593, 428/592, 428/607 , 428/192, 428/116, 428/298.1, 428/300.7, 428/297.4, 244/119, 244/133, 244/120, 244/126, 244/129.1, 244/117 R, 244/125
• US Class Current: 428/608
• CPC Class Codes: B32B 15/08 of synthetic resin
• CPC Class Codes: B32B 15/14 next to a fibrous or filamentary layer
• CPC Class Codes: B32B 2250/42 Alternating layers, e.g. ABAB(C), AABBAABB(C)
• CPC Class Codes: B32B 2255/06 on metal layer
• CPC Class Codes: B32B 2260/021 Fibrous or filamentary layer layers of natural mineral fibres or particles being impregnated in a plastic substance B32B19/02
• CPC Class Codes: B32B 2260/023 Two or more layers
• CPC Class Codes: B32B 2260/046 Synthetic resin
• CPC Class Codes: B32B 2262/10 Inorganic fibres layered products comprising a layer of natural mineral fibres B32B19/00
• CPC Class Codes: B32B 2262/106 Carbon fibres, e.g. graphite fibres
• CPC Class Codes: B32B 2305/076 Prepregs
• CPC Class Codes: B32B 2307/306 Resistant to heat
• CPC Class Codes: B32B 2307/54 Yield strength; Tensile strength
• CPC Class Codes: B32B 2307/552 Fatigue strength
• CPC Class Codes: B32B 2307/714 Inert, i.e. inert to chemical degradation, corrosion
• CPC Class Codes: B32B 2309/105 Thickness
• CPC Class Codes: B32B 2311/18 Titanium
• CPC Class Codes: B32B 2605/18 Aircraft blades, propellers B32B2603/00
• CPC Class Codes: B32B 3/12 characterised by a layer of regularly- arranged cells, e.g. a honeycomb structure
• CPC Class Codes: B32B 37/146 whereby one or more of the layers is a honeycomb structure
• CPC Class Codes: B32B 5/02 characterised by structural features of a fibrous or filamentary layer layer formed of metallic wires B32B15/02; layer formed of natural mineral fibres B32B19/02; layer formed of wood fibres B32B21/02
• 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 7/03 : with respect to the orientation of features direction of fibres or filaments B32B5/12
• CPC Class Codes: B64C 1/12 : Construction or attachment of skin panels
• CPC Class Codes: B64C 2001/0072 : from composite materials
• CPC Class Codes: B64C 3/26 : Construction, shape, or attachment of separate skins, e.g. panels
• CPC Class Codes: B64C 30/00 : Supersonic type aircraft
• CPC Class Codes: Y02T 50/40 : Weight reduction
• CPC Class Codes: Y10T 428/1234 : Honeycomb, or with grain orientation or elongated elements in defined angular relationship in respective components [e.g., parallel, inter- secting, etc.]
• CPC Class Codes: Y10T 428/12431 : Foil or filament smaller than 6 mils
• CPC Class Codes: Y10T 428/12444 : Embodying fibers interengaged or between layers [e.g., paper, etc.]
• CPC Class Codes: Y10T 428/12569 : Synthetic resin
• CPC Class Codes: Y10T 428/12806 : Refractory [Group IVB, VB, or VIB] metal-base component
• CPC Class Codes: Y10T 428/24165 : Hexagonally shaped cavities
• CPC Class Codes: Y10T 428/249942 : Fibers are aligned substantially parallel
• CPC Class Codes: Y10T 428/26 : Web or sheet containing structurally defined element or component, the element or component having a specified physical dimension
• CPC Class Codes: Y10T 428/31678 : Of metal

Abstract

The invention provides a hybrid laminate and skin panels of hybrid laminate structure that are suitable for a supersonic civilian aircraft. The hybrid laminates include layups of layers of titanium alloy foil and composite plies, that are optimally oriented to counteract forces encountered in use, that are bonded to a central core structure, such as titanium alloy honeycomb. The reinforcing fibers of the composite plies are selected from carbon and boron, and the fibers are continuous and parallel oriented within each ply. However, some plies may be oriented at angles to other plies. Nevertheless, in a preferred embodiment of the invention, a substantial majority of, or all of, the fibers of the hybrid laminates are oriented in a common direction. The outer surfaces of the laminates include a layer of titanium foil to protect the underlying composite-containing structure from the environment, and attack by solvents, and the like.

Claims

1. A composite aircraft fuselage section comprising a symmetrical hybrid laminate layup, the layup comprising laminated together:

(a) a first layer of metal foil comprising an outer surface of the fuselage section;

(b) a second layer of metal foil comprising another layer of the laminate layup; and

(c) at least one hoop ply interposed between the first and second layers, and bonded to at least one of the first and second layers, the at least one hoop ply comprising commonly aligned fibers embedded in a polymeric matrix, wherein the at least one hoop ply has fibers extending helically around the circumference of the aircraft fuselage section.

2. The fuselage section of claim 1, wherein the at least one hoop ply comprises two hoop plies, a first of the two hoop plies bonded to an inboard surface of the first layer of foil, and a second of the two hoop plies bonded to an outboard surface of the second layer of foil, wherein inboard is the direction towards a longitudinal central axis of the aircraft fuselage section, and outboard is the direction away from a longitudinal axis of the aircraft fuselage section.

3. The fuselage section of claim 2, further comprising at least one composite layer having longitudinal oriented fibers, the at least one composite layer interposed between the first and second hoop plies.

4. The fuselage section of claim 3, wherein the at least one composite layer comprises three plies.

5. The fuselage section of claim 2, further comprising a central layer of metal foil interposed between the first and second metal foil layers.

6. The fuselage section of claim 5, further comprising at least three longitudinal plies, defined as plies having longitudinal oriented fibers, interposed between the central layer of metal foil and each of the two hoop plies.

7. The fuselage of claim 2, further comprising a third layer of metal foil bonded to an inboard surface of the first hoop ply;

a fourth layer of metal foil bonded to an outboard surface of the second hoop ply; and

a layer of longitudinal plies interposed between, and bonded to, both the third and fourth metal foil layers, wherein longitudinal plies are defined as plies having longitudinal oriented fibers.

8. The fuselage section of claim 2, including:

(a) a crown portion comprising a central layer of metal foil interposed between the first and second metal foil layers and two longitudinal plies, one longitudinal ply interposed between the central layer of metal foil and each of the two hoop plies, and (b) a side portion comprising a third layer of metal foil bonded to the inboard surface of the first hoop ply;

a fourth layer of metal foil bonded to the outboard surface of the second hoop ply; and

a longitudinal ply interposed between, and bonded to, both the third and fourth metal foil layers, wherein the longitudinal plies are defined as plies having longitudinal oriented fibers.

9. The fuselage section of claim 6, wherein the at least three longitudinal plies interposed between the central layer of metal foil and each of the two hoop plies comprise boron fibers.

10. The fuselage section of claim 2, including:

(a) a crown portion comprising a central layer of metal foil interposed between the first and second metal foil layers and two longitudinal plies, defined as plies having longitudinal oriented fibers, one longitudinal ply interposed between the central layer of metal foil and each of the two hoop plies, and (b) a keel portion comprising two central layers of metal foil interposed between the first and second foil layers and three longitudinal plies, defined as plies having longitudinal fibers, the first longitudinal ply interposed between the two central layers of metal foil, the second longitudinal ply interposed between the one of said two central layers of metal foil closest to said first hoop ply and the first hoop ply and the third longitudinal ply interposed between the other of said two central layers of metal foil and the second hoop ply, each of the longitudinal plies including boron fibers.

11. The fuselage section of claim 1, wherein the at least one ply is from about 0.005 to about 0.03 inches thick.

12. The fuselage section of claim 1, wherein the first and second layers of metal foil comprise a titanium alloy.

13. The fuselage section of claim 1, wherein the first and second layers of metal foil comprise a heat treated beta titanium alloy.

14. The fuselage section of claim 1, wherein the first and second layers of metal foil are each of thickness in the range from about 0.003 to about 0.01 inches.

15. The fuselage section of claim 1, wherein the first and second layers of metal foil comprise butt-joined foils.

16. The fuselage section of claim 1, wherein the matrix is resistant to repeated exposure to temperatures of at least about 350°

F.

17. A composite aircraft fuselage section comprising a central honeycomb core layer, and a hybrid layup bonded to each side of the central honeycomb core layer to form a hybrid laminate layup that is symmetrical about the central honeycomb core layer, each of the hybrid layups comprising laminated together:

(a) first layer of metal foil comprising an outer surface of the fuselage section;

(b) a second layer of metal foil comprising another layer of the laminate layup; and

(c) at least one hoop ply interposed between the first and second layers, and bonded to at least one of the first and second layers the at least one hoop ply comprising commonly aligned fibers embedded in a polymeric matrix, wherein the at least one hoop ply has fibers extending helically around the circumference of the aircraft fuselage section.

18. A composite aircraft fuselage section, the fuselage section comprising a symmetrical hybrid laminate, the laminate comprising:

(a) a pair of layups, each of the layups comprising;

(i) a heat treated beta titanium alloy foil layer comprising butt-joined foils each of thickness in the range from about 0.003 to about 0.01 inches; and

(ii) a layer of polymeric composite bonded to a side of the foil layer, the polymeric layer comprising at least one ply comprising a polymeric matrix, the matrix being resistant to repeated exposure to temperatures of at least about 350°

F., and the composite having commonly aligned fibers embedded in the matrix; and

(b) a central honeycomb core layer, each of the pair of layups bonded to one side of the core layer to form the symmetrical hybrid laminate.

19. The fuselage section of claim 18, wherein the polymer is selected from the group consisting of polyaryletherketone, polyetheretherketone, polyimides, polyarylethersulfone, oxydiphthalic dianhydride 3, 4'"'"' oxydianiline, and functional derivatives thereof.

20. The fuselage section of claim 18, wherein the at least one ply is from about 0.005 to about 0.03 inches thick.

21. The fuselage section of claim 18, wherein the foil layer is heat-treated to a yield strain of greater than 1%.

22. The fuselage section of claim 18, wherein the foil layer is pretreated to produce a surface for more tenacious bonding to the polymeric matrix.

23. The fuselage section of claim 18, wherein the fibers are selected from the group consisting of carbon and boron fibers.

24. The fuselage section of claim 18, wherein the fibers are continuous fibers.

25. The fuselage section of claim 18, wherein the open-hole tensile strength of the laminate is greater than about 55% of the unnotched ultimate strength of the laminate.

26. The fuselage section of claim 18, wherein the open-hole compression strength of the laminate is at least about 50 ksi.

27. The fuselage section of claim 18, wherein crack-growth rate, after crack initiation in the hybrid laminate, is less than about 0.2% of the crack-growth rate of the titanium alloy in monolithic form.

28. The fuselage section of claim 18, wherein a majority of the fibers of the hybrid laminate are aligned in a common direction.

29. The fuselage section of claim 18, further comprising another layer of polymeric composite, the another layer comprising commonly aligned reinforcing fibers, the fibers aligned at 90°

relative to fibers of the layer of the polymeric composite of (a)(ii).

30. The fuselage section of claim 29, further comprising another titanium alloy foil layer, the another foil layer bonded to a side of the another layer of the polymeric composite.

31. The fuselage section of claim 18, further comprising additional layers of polymeric composite, the additional layers comprising commonly aligned reinforcing fibers, the fibers of at least one of the addition layers aligned at 90°

relative to fibers of the layer of the polymeric composite of (a)(ii).

32. The fuselage section of claim 31, further comprising at least one addition titanium alloy foil layer, the at least one addition titanium alloy foil layer bonded to a side of at least one of the additional layers of the polymeric composite.

33. The fuselage section of claim 18, wherein the at least one ply of each composite layer comprises four plies.

34. The fuselage section of claim 18, wherein each of the layups further comprises a plurality of composite layers and titanium alloy foil layers, wherein a titanium foil layer is interposed between each pair of adjacent composite layers.

35. The fuselage section of claim 18, wherein each of the pair of layups comprises a first and second layer of metal foil, the layers of metal foil separated by an intervening layer of polymeric composite bonded to sides of each of the layers of foil, and wherein each of the layers of foil comprise butt joined gore panels such that butt joints of the first layer are offset from butt joints of the second layer.

36. The fuselage section of claim 35, wherein the first layer of metal foil of each of the layups is an outer later of the hybrid laminate.

37. A composite aircraft fuselage section, the fuselage section comprising a symmetrical hybrid laminate, the laminate comprising:

(a) a honeycomb core having two sides; and

(b) a pair of layups, one of the pair bonded to each side of the core, each of the layups comprising laminated layers of;

(i) heat treated beta titanium alloy gore panels having a thickness in the range from about 0.003 to about 0.01 inches, the panels butt-joined to form a continuous foil layer; and

(ii) an organic polymeric composite layer between adjacent foil layers, the composite layer comprising at least one ply, the at least one ply comprising a polymeric matrix resistant to repeated exposure to high temperatures encountered in supersonic flight, the polymeric matrix having commonly aligned reinforcing fibers embedded therein.

38. The aircraft fuselage section of claim 37, wherein each of the layups further comprises an outer layer of titanium alloy foil, the outer layer bonded to an adjacent composite ply.

39. The aircraft fuselage section of claim 37, wherein the high temperature resistant matrix is selected from the group consisting of polyaryletherketone, polyetheretherketone, polyimides, polyarylethersulfone, oxydiphthalic dianhydride 3, 4'"'"' oxydianiline, and functional derivatives thereof.

40. The aircraft fuselage section of claim 37, wherein the foil layers have a thickness of from about 0.003 to about 0.01 inches.

41. The aircraft fuselage section of claim 37, wherein the open-hole tensile strength of each of the layups is greater than about 55% of the unnotched ultimate strength of the layup.

42. The aircraft fuselage section of claim 37, wherein the open-hole compressive strength of each of the layups is at least about 50 ksi.

43. The aircraft fuselage section of claim 37, wherein the crack-growth rate, after crack initiation in the aircraft skin panel, is less than about 0.2% of the crack-growth rate of the titanium alloy in monolithic form.

44. The aircraft fuselage section of claim 37, wherein the reinforcing fibers are selected from the group consisting of graphite and boron fibers.

45. The aircraft fuselage section of claim 37, wherein the ultimate tensile strength exceeds about 2×

10⁶ psi/lb/in³.

46. The aircraft fuselage section of claim 37, wherein the ultimate compressive strength exceeds about 1.5×

10⁶ psi/lb/in³.

47. The aircraft fuselage section of claim 37, wherein each layup comprises a first and second layer of metal foil, the layers of metal foil separated by an intervening layer of polymeric composite butt-joined gore panels such that butt joints of the first layer are offset from butt joints of the second layer.

48. The aircraft fuselage section of claim 37, wherein the first layer of metal foil of each of the layups is an outer layer of the hybrid laminate.