Titanium-polymer hybrid laminates

• US 5,866,272 A
• Filed: 01/11/1996
• Issued: 02/02/1999
• 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): Zimmerman, John J.
• Assistant Examiner(s): LA VILLA, MICHAEL EUGENE
• Application Number:
  US08/585,304
• Time in Patent Office: 1,118 Days
• Field of Search: 428/593, 428/606, 428/608, 428/615, 428/292, 428/293, 428/458, 428/660, 428/220, 428/332, 428/457, 428/626, 428/118, 244/119
• US Class Current: 428/593
• 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 symmetrical hybrid 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 resistant to repeated exposure to temperatures of at least about 350°

F., 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.

2. The hybrid laminate of claim 1, wherein the polymer is selected from the group consisting of polyaryletherketone, polyetheretherketone, polyimides, polyarylethersulfone, oxydiphthalic dianhydride 3, 4'"'"' oxydianiline, and functional derivatives thereof.

3. The hybrid laminate of claim 1, wherein the at least one ply is from about 0.005 to about 0.03 inches thick.

4. The hybrid laminate of claim 1, wherein the foil layer is heat-treated to a yield strain of greater than 1%.

5. The hybrid laminate of claim 1, wherein the foil layer is pretreated to produce a surface for more tenacious bonding to the polymeric matrix.

6. The hybrid laminate of claim 1, wherein the fibers are selected from the group consisting of carbon and boron fibers.

7. The hybrid laminate of claim 1, wherein the fibers are continuous fibers.

8. The hybrid laminate of claim 1, wherein the open-hole tensile strength of the laminate, is greater than about 55% of the unnotched ultimate strength of the laminate.

9. The hybrid laminate of claim 1, wherein the open-hole compression strength of the laminate is at least about 50 ksi.

10. The hybrid laminate of claim 1, wherein crack-growth rate, after crack initiation in the hybrid laminate, is less than about 0.2% of the crack-grow rate of the titanium alloy in monolithic form.

11. The hybrid laminate of claim 1, wherein a majority of the fibers of the hybrid laminate are aligned in a common direction.

12. The hybrid laminate of claim 1, wherein each of the pair of layups comprises a first and a 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.

13. The hybrid laminate of claim 12, wherein the first layer of metal foil of each of the layups is an outer layer of the hybrid laminate.

14. The hybrid laminate of claim 1, 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 polymeric composite of (a)(ii).

15. An aircraft skin panel, the panel comprised of 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.

16. The aircraft skin panel of claim 15, wherein each of the layups further comprises an outer layer of titanium alloy foil, the outer layer bonded to an adjacent composite ply.

17. The aircraft skin panel of claim 15, 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.

18. The aircraft skin panel of claim 15, wherein the open-hole tensile strength of each of the layups is greater than about 55% of the unnotched ultimate strength of the layup.

19. The aircraft skin panel of claim 15, wherein the open-hole compressive strength of each of the layups is at least about 50 ksi.

20. The aircraft skin panel of claim 15, 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.

21. The aircraft skin panel of claim 15, wherein the reinforcing fibers are selected from the group consisting of graphite and boron fibers.

22. The aircraft skin panel of claim 15, wherein the ultimate tensile strength exceeds about 2×

10⁶ psi/lb/in³.

23. The aircraft skin panel of claim 15, wherein the ultimate compressive strength exceeds about 1.5×

10⁶ psi/lb/in³.

24. The aircraft skin panel of claim 15, wherein each layup comprises a first and a 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.

25. The aircraft skin panel of claim 24, wherein the first layer of metal foil of each of the layups is an outer layer of the hybrid laminate.