Light weight and high strength fuselage

• US 6,648,273 B2
• Filed: 10/30/2001
• Issued: 11/18/2003
• Est. Priority Date: 10/30/2001
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Anast, Peter Z.
• Primary Examiner(s): Poon, Peter M.
• Assistant Examiner(s): COLLINS, TIMOTHY D
• Application Number:
  US10/017,732
• Publication Number:
  US 20030080251A1
• Time in Patent Office: 749 Days
• Field of Search: 244/117.R, 244/119, 244/120, 244/121, 244/123, 244/133, 244/132
• US Class Current: 244/119
• CPC Class Codes: B64C 1/068 Fuselage sections
• CPC Class Codes: B64C 1/12 Construction or attachment of skin panels
• CPC Class Codes: B64C 2001/0081 from metallic materials
• CPC Class Codes: Y02T 50/40 Weight reduction

Abstract

A fuselage comprising a skin assembly including an outer, laminate skin bonded to an inner, aluminum doubler. The fuselage also includes a support structure comprising a plurality of longitudinal stringer members and a plurality of annular frame members that are attached to, and cooperate to support, the skin assembly. The aluminum doubler provides additional structural support for the fuselage, and in particular, for the outer laminate skin of the skin assembly. The additional structural strength added by the aluminum doubler allows the use of an improved range of fasteners, such as knife-edge, countersink rivets and further allows the use of the laminate layer even in areas with a large number of cutouts, such as the window track of the fuselage. The members of the support structure may interconnected via a plurality of integral flanges, which, when combined with the skin, provide improved structural strength for the entire fuselage.

Claims

1. A fuselage assembly for an aerospace craft defining a long axis, said fuselage assembly comprising:

a plurality of longitudinal stringer members radially spaced from, and extending generally parallel to, the longitudinal axis of the fuselage, said longitudinal stringer members further spaced circumferentially from each other, each of the longitudinal stringer members having a stringer wall structure including an outer longitudinal surface;

a plurality of annular frame members spaced along the longitudinal axis, each of the annular frame members having a frame wall structure including a plurality of outer circumferential surfaces, each of the outer circumferential surfaces interposed between a respective pair of the longitudinal stringer members;

a lightweight doubler attached to and covering at least a portion of the outer surfaces of the frame and stringer members; and

a laminate sheet comprising alternating layers of metal and composite, said laminate sheet disposed over, and attached to, the lightweight doubler so as to form an outer skin of the fuselage strengthened by the underlying lightweight doubler, the annular frame members and the stringer members, wherein at least a portion of the laminate sheet disposed over the lightweight doubler defines an opening therethrough and has a constant gauge thickness approximately equal to another, remote portion of the laminate sheet that is free of openings and not disposed over the lightweight doubler.

2. A fuselage assembly of claim 1, wherein the stringer wall structure of each of the longitudinal stringers further includes a flange defining the outer longitudinal surface.

3. A fuselage assembly of claim 2, wherein the frame wall structure of each of the frame members includes a plurality of flanges, each of the flanges defining a respective one of the outer circumferential surfaces.

4. A fuselage assembly of claim 3, wherein each of the flanges of the frame wall structure overlaps a portion of the flange of each of the respective pair of longitudinal stringer members.

5. A fuselage assembly of claim 4, wherein the overlapping flange portions, the lightweight doubler and the laminate sheet are attached together using a fastener that extends through the opening in the constant gauge laminate sheet portion of the outer skin.

6. A fuselage assembly of claim 5, wherein the fastener is a knife-edge, countersink fastener.

7. A fuselage assembly of claim 1, wherein the laminate sheet is bonded to the lightweight doubler using an adhesive layer.

8. A fuselage assembly of claim 1, wherein the lightweight doubler is constructed of aluminum.

9. A fuselage assembly of claim 1, wherein the metal layers of the laminate sheet are aluminum layers and the composite layers are fiberglass-epoxy layers.

10. A fuselage assembly of claim 1, wherein the opening in the constant gauge laminate sheet portion of the outer skin defines one of a plurality of cutouts.

11. A fuselage assembly of claim 10, wherein each of the plurality of cutouts houses one of a plurality of windows.

12. A method of assembling a skin of a lightweight fuselage to a support structure including a plurality of longitudinal stinger members and annular frame members, the method of attaching the skin comprising:

positioning the stringer members and frame members so that the members define a plurality of outer surfaces;

bonding a plurality of lightweight doubler panels to respective laminate skin panels to form a plurality of skin assembly panels;

positioning each of the skin assembly panels over a respective portion of the plurality of outer surfaces; and

attaching each of the skin assembly panels to the respective portion of the plurality of outer surfaces by transfixing the skin assembly panels and the members with a plurality of fasteners wherein at least one of the fasteners is transfixed through a portion of one of the skin assembly panels including the laminate skin and doubler, said laminate skin having a constant gauge thickness with respect to another remote portion of the laminate skin not disposed over the lightweight doubler.

13. A method of claim 12, wherein bonding includes anodizing surfaces of the doubler panels and laminate skin panels.

14. A method of claim 13, wherein bonding includes application of an adhesive primer to the surfaces of the doubler panels and laminate skin panels.

15. A method of claim 12, wherein the plurality of fasteners are a plurality of knife-edged rivets and wherein attaching each of the panels includes countersinking the knife-edged rivets.

16. A method of claim 12, wherein positioning includes splicing the frame members with the longitudinal members.

17. A method of claim 12, wherein positioning includes forming a plurality of continuous circumferential outer surfaces with the members and wherein attaching the skin assembly panels includes attaching the doubler panels adjacent the circumferential outer surfaces and arresting skin crack propagation.