Splice joints for composite aircraft fuselages and other structures

• US 8,061,035 B2
• Filed: 01/18/2008
• Issued: 11/22/2011
• Est. Priority Date: 09/23/2004
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Rolfes, Neal G., Clapp, Brian C., Chan, Wallace C., Stulc, Jeffrey F.
• Primary Examiner(s): Bryant; David
• Assistant Examiner(s): Walters; Ryan J
• Application Number:
  US12/016,258
• Publication Number:
  US 20080111026A1
• Time in Patent Office: 1,404 Days
• Field of Search: 298/97, 298/972, 294/62, 294/63, 294/64, 294/68, 527/831.4, 527/831.5, 527/871.2, 528/001.1, 244/2, 244/117.R, 244/119, 244/120, 244/131, 244/132, 244/133
• US Class Current: 29/897.2
• CPC Class Codes: B64C 1/069 Joining arrangements therefor
• CPC Class Codes: B64C 1/12 Construction or attachment of skin panels
• CPC Class Codes: B64C 1/1492 Structure and mounting of the transparent elements in the window or windscreen
• CPC Class Codes: B64C 2001/0072 from composite materials
• CPC Class Codes: Y02T 50/40 Weight reduction
• CPC Class Codes: Y10T 29/49616 Structural member making
• CPC Class Codes: Y10T 29/49622 Vehicular structural member making
• CPC Class Codes: Y10T 29/49826 Assembling or joining
• CPC Class Codes: Y10T 428/19 Sheets or webs edge spliced or joined
• CPC Class Codes: Y10T 428/24331 including nonapertured component
• CPC Class Codes: Y10T 428/24612 Composite web or sheet
• CPC Class Codes: Y10T 428/24628 Nonplanar uniform thickness material
• CPC Class Codes: Y10T 428/24661 Forming, or cooperating to form cells

Abstract

Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Claims

1. A method for manufacturing a shell structure, the method comprising:

positioning a first composite skin in edgewise alignment with a second composite skin;

attaching a first end of a fitting to the first composite skin and a first stiffener on the first composite skin, wherein the first stiffener has a first flange portion attached to the first composite skin and a first raised portion projecting away from the first composite skin; and

attaching a second end of the fitting to the second composite skin and a second stiffener on the second composite skin to join the first composite skin to the second composite skin, wherein the second stiffener has a second flange portion attached to the second composite skin and a second raised portion projecting away from the second composite skin, wherein attaching the first end portion of the fitting to the first stiffener includes attaching the first end portion to the first flange portion,

and wherein attaching the second end portion of the fitting to the second stiffener includes attaching the second end portion to the second flange portion.

2. The method of claim 1, further comprising attaching a strap to a first edge region of the first composite skin and a second edge region of the second composite skin to splice the first and second composite skins together before the fitting is attached.

3. The method of claim 1, further comprising attaching a strap to a first edge region of the first composite skin and a second edge region of the second composite skin to splice the first and second composite skins together before the fitting is attached, and wherein attaching the fitting includes sandwiching a portion of the strap between the fitting and the first edge region of the first composite skin, and sandwiching a second portion of the strap between the fitting and the second edge region of the second composite skin.

4. The method of claim 1 wherein the first composite skin extends 360 degrees around a first axis, and wherein the second composite skin extends 360 degrees around a second axis.

5. The method of claim 1 wherein attaching a first end of a fitting to the first stiffener and a second end of the fitting to the second stiffener includes leaving end portions of the first and second stiffeners open for liquid drainage.

6. A method for manufacturing an aircraft fuselage, the method comprising:

attaching at least a first stiffener to a first composite skin, wherein the first stiffener has a first flange portion attached to the first composite skin and a first raised portion projecting away from the first composite skin, the first composite skin extending 360 degrees around a first axis;

attaching at least a second stiffener to a second composite skin, wherein the second stiffener has a second flange portion attached to the second composite skin and a second raised portion projecting away from the second composite skin, the second composite skin extending 360 degrees around a second axis;

positioning the first composite skin adjacent to the second composite skin so that the first and second axes are at least approximately aligned; and

attaching a first end portion of a fitting to the first stiffener and the first composite skin and a second end portion of the fitting to the second stiffener and the second composite skin to join the first composite skin to the second composite skin, wherein attaching the first end portion of the fitting to the first stiffener includes attaching the first end portion to the first flange portion, and wherein attaching the second end portion of the fitting to the second stiffener includes attaching the second end portion to the second flange portion.

7. The method of claim 6, further comprising:

attaching at least a third stiffener to the first composite skin;

attaching at least a fourth stiffener to the second composite skin; and

attaching the first end portion of the fitting to the third stiffener and the second end portion of the fitting to the fourth stiffener.

8. The method of claim 6, further comprising:

attaching at least a third stiffener to the first composite skin so that a first flange portion of the first stiffener and a third flange portion of the third stiffener form an at least approximately continuous first support surface;

attaching at least a fourth stiffener to the second composite skin so that the second flange portion of the second stiffener and a fourth flange portion of the fourth stiffener form an at least approximately continuous second support surface, wherein attaching a first end portion of a fitting to the first stiffener includes attaching the first end portion to the first support surface, and wherein attaching a second end portion of the fitting to the second stiffener includes attaching the second end portion to the second support surface.

9. The method of claim 6 wherein positioning the first composite skin adjacent to the second composite skin includes positioning a first barrel section having a first width of about 15 feet or more with a second barrel section having a second width of about 15 feet or more.

10. The method of claim 6, further comprising:

forming a first portion of a window cut-out in a first edge region of the first composite skin; and

forming a second portion of a window cut-out in a second edge region of the second composite skin, wherein positioning the first composite skin adjacent to the second composite skin includes aligning the first portion of the window cut-out with the second portion of the window cut-out prior to joining the first composite skin to the second composite skin.

11. The method of claim 10, further comprising attaching a strap to the first edge region of the first skin and the second edge region of the second skin to splice the first and second composite skins together, wherein the strap includes an aperture at least approximately aligned with the first and second portions of the window cut-out.

12. The method of claim 6, further comprising attaching a strap to a first edge region of the first composite skin and a second edge region of the second composite skin to splice the first and second composite skins together before attaching the first end portion of the fitting to the first stiffener and the first composite skin and the second end portion of the fitting to the second stiffener and the second composite skin.

13. The method of claim 6, further comprising attaching a strap to a first edge region of the first composite skin and a second edge region of the second composite skin to splice the first and second composite skins together, and wherein attaching the fitting includes sandwiching a portion of the strap between the fitting and the first edge region of the first composite skin, and sandwiching a second portion of the strap between the fitting and the second edge region of the second composite skin.

14. The method of claim 6, further comprising:

attaching a third flange portion of a third stiffener to the first composite skin;

attaching a fourth flange portion of a fourth stiffener to the second composite skin;

attaching the first end portion of the fitting to the third flange portion of the third stiffener; and

attaching the second end portion of the fitting to the fourth flange portion of the fourth stiffener.

15. The method of claim 6, further comprising forming the fitting from composite materials.

16. The method of claim 6, further comprising:

forming the first and second stiffeners from composite materials; and

forming the fitting from composite materials.