Composite Aircraft Structures With Hat Stiffeners

• US 20080111024A1
• Filed: 11/10/2006
• Published: 05/15/2008
• Est. Priority Date: 11/10/2006
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Lee, Shin S., Viisoreanu, Adrian, Koch, William J.
• Granted Patent: US 7,871,040 B2
• Time in Patent Office: Days
• US Class Current: 244/121
• CPC Class Codes: B64C 1/068 Fuselage sections
• CPC Class Codes: B64C 1/12 Construction or attachment of skin panels
• CPC Class Codes: B64C 2001/0072 from composite materials
• CPC Class Codes: Y02T 50/40 Weight reduction

Abstract

A composite panel structure for an aircraft includes a plurality of annular, hat-shaped frames disposed coaxially along a long axis of the aircraft in a spaced, parallel relationship, an inner skin having an inner surface bonded to an outer surface of the hat frames, a plurality of elongated, hat-shaped stringers disposed in a longitudinal direction along an outer surface of the inner skin in a spaced, parallel relationship, a solid or rigid foam offset bonded to an outer surface of each of the stringers, and an outer skin having an inner surface bonded to an upper surface of each of the offsets. The inner skin carries the loads of the structure and the outer skin defines an aerodynamic surface of the aircraft and provides impact and lighting protection. The frames, inner skin and stringers are formed on a single forming tool and cured and bonded with each other simultaneously.

Claims

1. A composite panel structure for an aircraft, comprising:

a plurality of annular frames disposed coaxially along a long axis of the aircraft in a spaced, parallel relationship, each frame having a hat shaped cross section;

an inner skin having an inner surface bonded to an outer surface of the hat frames;

a plurality of elongated stringers disposed in a longitudinal direction along an outer surface of the inner skin in a spaced, parallel relationship, each stringer having a hat shaped cross section;

an offset bonded to an outer surface of each of the stringers; and

,

an outer skin having an inner surface bonded to an upper surface of each of the offsets.

2. The composite panel structure of claim 1, wherein each frame comprises:

a circumferential inner chord;

a pair of circumferential risers, each extending at an angle from a respective opposite side of the inner chord; and

,

a pair of circumferential flanges, each extending longitudinally from an outer end of a respective riser and in a direction opposite to the other flange, and

wherein the inner surface of the inner skin is bonded to an outer surface of each of the circumferential flanges of each frame.

3. The composite panel structure of claim 1, wherein each stringer comprises:

a longitudinal outer chord;

a pair of longitudinal risers, each extending at an angle from a respective opposite side of the outer chord; and

,

a pair of longitudinal flanges, each extending circumferentially from an inner end of a respective riser and in a direction opposite to the other flange, and

wherein the offsets are respectively bonded to an outer surface of the outer chord of a respective stringer.

4. The composite panel structure of claim 1, wherein the inner skin functions to carry substantially the entire design load of the structure.

5. The composite panel structure of claim 1, wherein the outer skin functions as the aerodynamic surface of the aircraft, to provide impact resistance, and to protect the aircraft against lightning strikes.

6. The composite panel structure of claim 1, wherein each of the offsets comprises a rigid, closed-cell foam material, or a solid high loss factor polymer material.

7. The composite panel structure of claim 1, wherein the offsets are bonded to the stringers and the outer skin is bonded to the offsets with an adhesive.

8. The composite panel structure of claim 1, wherein each of the offsets comprises a uniform thin layer or incorporates a web-like pattern.

9. The composite panel structure of claim 1, wherein the offsets further comprise columns disposed between the stringers and having opposite ends, at least one of which is bonded to an adjacent one of the inner and the outer skins with an adhesive.

10. The composite panel structure of claim 1, wherein longitudinal flanges of the stringers overlap circumferential flanges of the frames on opposite sides of the inner skin at a plurality of locations thereon, and further comprising a fastener extending through the overlapping flanges and the inner skin at each such location.

11. The composite panel structure of claim 1, further comprising a layer of fiberglass insulation disposed between the inner and outer skins and adjacent pairs of the stringers.

12. The composite panel structure of claim 1, wherein each of the frames comprises alternating bands of zero and ±

45 degree prepreg tape.

13. The composite panel structure of claim 1, wherein each of the frames comprises alternating bands of zero and ±

45 degree prepreg tape.

14. An aircraft fuselage incorporating the composite panel structure of claim 1.

15. A method for repairing damage to the outer surface of the outer skin of the aircraft of claim 14, the method comprising:

removing the damaged portion of the outer surface of the skin; and

,

cold-bonding a patch to the skin in place of the portion removed.

16. A method for making a composite panel structure for an aircraft, the method comprising:

providing a segmented internal mold line (IML) tool containing a plurality of spaced, parallel circumferential grooves, each adapted to form a respective one of a plurality of annular hat frames therein;

winding a plurality of layers of a prepreg tape in each groove of the tool to form an uncured hat frame therein;

placing a mandrel in a cavity of each of the hat frames;

winding a plurality of layers of a prepreg tape over the tool, hat frames and mandrels to form an uncured composite inner skin that is in contact with circumferential flanges of the hat frames at a co-cured bond line;

locating a plurality spaced, parallel, longitudinally extending mandrels for forming hat stringers on an outer surface of the uncured inner skin;

placing a plurality of layers of a prepreg tape over each of the hat stringer forming mandrels to form an uncured hat stringer thereon, each hat stringer having a pair of oppositely directed longitudinal flanges thereon, each of which is in contact with the outer surface of the inner skin at a co-cured bond line; and

,

heating the tool and the assembly of uncured hat frames, inner skin and hat stringers to cure a resin in the respective prepreg tapes thereof and to bond the respective flanges of the hat frames and stiffeners to the inner skin.

17. The method of claim 16, wherein the longitudinal flanges of the hat stringers overlap the circumferential flanges of the hat frames on opposite sides of the inner skin at a plurality of locations thereon, and further comprising installing a fastener through the overlapping flanges and the inner skin at each such location.

18. The method of claim 16, further comprising bonding an offset on an upper surface of an upper chord of each of the hat stringers.

19. The method of claim 18, further comprising bonding an inner surface of an outer skin to an upper surface of each of the offsets.

20. The method of claim 16, wherein the mandrels placed in the respective cavities of the hat frames comprise fly-away mandrels that are left in the structure after it is cured.

21. The method of claim 16, wherein locating the hat stringer forming mandrels comprises sliding the mandrels over respective longitudinal rows of radial locator pins disposed around a perimeter of the tool.

22. The method of claim 16, wherein the hat stringer forming mandrels comprise aluminum and are removed from the structure after it is cured.

23. An aircraft having a fuselage made in accordance with the method of claim 16.