Bonded and tailorable composite assembly

Patent metadata

Field	Value
Publication number	US-10155581-B2
Application number	US-201615191527-A
Country	US
Kind code	B2
Filing date	Jun 24, 2016
Priority date	Aug 28, 2012
Publication date	Dec 18, 2018
Grant date	Dec 18, 2018

Abstract

An all-composite assembly such as a composite laminate aircraft empennage has vertical and horizontal stabilizers with differing sets of interlaminar fracture toughnesses and differing stiffnesses to improve flight characteristics. Composite laminate skins are bonded to unitized and stiffened understructure to reduce weight and improve damage containment.

First claim

What is claimed is:

1. A method of making a composite structure, comprising:
fabricating a plurality of composite spars;
fabricating a plurality of composite stiffeners;
forming a composite understructure by bonding the composite spars and the composite stiffeners together, such that:
the composite structure comprises the composite understructure and a composite first laminate skin bonded to the understructure;
the composite first laminate skin comprises a first set of pre-selected interlaminar fracture toughnesses; and
the understructure comprises:
a plurality of longitudinally extending composite spars, and
a plurality of Z-shaped composite stiffeners extending between and bonded to the spars, and a plurality of longitudinally extending, composite stringers passing through the Z-shaped composite stiffeners and bonded to at least one of: the composite first laminate skin, and composite second laminate skin;

at least one second composite structure, the second composite structure comprising a composite second understructure and a composite laminate second skin bonded to the second understructure, the composite laminate second skin comprising a second set of pre-selected interlaminar fracture toughnesses; and
bonding first and second composite skins on opposite sides of the composite understructure.

2. The method of claim 1 , further comprising:
each of the composite first laminate skin and the composite laminate second skin being subject to Mode I, II and III loading; and
the first set of pre-selected interlaminar fracture toughness and the second set of pre-selected interlaminar fracture toughnesses differing from each other in the Modes I, II, and III loading.

3. The method of claim 1 , wherein:
the composite structure comprises a first torsional stiffness; and
the second composite structure comprises a second torsional stiffness, the first torsional stiffness being greater than the second torsional stiffness.

4. The method of claim 3 , wherein:
the first torsional stiffness comprises a range comprising approximately 45.0 to 52.0 million pounds per square inch; and
the second torsional stiffness comprises a range comprising approximately 40.0 to 50.2 million pounds per square inch.

5. The method of claim 1 , wherein:
the first set of pre-selected interlaminar fracture toughnesses of the composite laminate first skin of the composite structure comprises:
a Mode I interlaminar fracture toughness within a range that comprises approximately 4.0 to 6.5 inch-pounds per square inch;
a Mode II interlaminar fracture toughness within a range that comprises approximately 12.0 to 15.5 inch-pounds per square inch; and
a Mode III interlaminar fracture toughness within a range that comprises approximately 16.0 to 18.5 inch-pounds per square inch.

6. The method of claim 5 , wherein the second set of pre-selected interlaminar fracture toughnesses of the composite laminate second skin of the second composite structure comprises:
a Mode I interlaminar fracture toughness within a range that comprises approximately 2.5 to 3.5 inch-pounds per square inch;
a Mode II interlaminar fracture toughness within a range that comprises approximately 7.5 to 9.5 inch-pounds per square inch; and
a Mode III interlaminar fracture toughness within a range that comprises approximately 18.0 to 20.5 inch-pounds per square inch.

7. The method of claim 1 , further comprising:
passing a plurality of substantially straight composite cross-beams respectively through the Z-shaped composite stiffeners and extending substantially normal to the composite spars.

8. The method of claim 1 , further comprising:
forming an aircraft empennage via arranging the composite structure and the second composite structure.

9. The method of claim 1 , further comprising:
an aircraft vertical stabilizer that comprises the composite structure; and
an aircraft horizontal stabilizer that comprises the second composite structure.

10. A method of forming an empennage for an aircraft, the method comprising:
forming composite first understructure via forming a first integrated grid via bonding together: first composite spars, first composite cross-beams, and first composite stiffeners;
forming a vertical stabilizer via bonding the composite first understructure to a composite laminate first skin, wherein the composite laminate first skin comprises:
a Mode I interlaminar fracture toughness within a range comprising approximately 4.0 to 6.5 inch-pounds per square inch;
a Mode II interlaminar fracture toughness within a range comprising approximately 12.0 to 15.5 inch-pounds per square inch; and
a Mode III interlaminar fracture toughness within a range comprising approximately 16.0 to 18.5 inch-pounds per square inch; and

forming a pair of horizontal stabilizers, via forming each of the horizontal stabilizers respectively via bonding comprising a composite second understructure a composite laminate second skin, the composite second understructure comprising a second integrated grid that comprises second composite spars, second composite cross-beams, and second composite stiffeners bonded together.

11. The method of claim 10 , wherein the composite laminate second skin comprises:
a Mode I interlaminar fracture toughness within a range that comprises approximately 2.5 to 3.5 inch-pounds per square inch;
a Mode II interlaminar fracture toughness within a range that comprises approximately 7.5 to 9.5 inch-pounds per square inch; and
a Mode III interlaminar fracture toughness within a range that comprises approximately 18.0 to 20.5 inch-pounds per square inch.

12. The method of claim 10 , wherein the vertical stabilizer comprises a torsional stiffness in a range that comprises approximately 45.0 to 52.0 million pounds per square inch.

13. The method of claim 10 , wherein each of the horizontal stabilizers comprises a bending stiffness in a range that comprises approximately 30.0 to 36.5 million pounds per square inch.

14. The method of claim 10 , wherein each of the spars comprises a bending stiffness of approximately 45 million pounds per square inch.

15. The method of claim 10 , wherein:
each of the composite stiffeners comprises a Z-shape; and
each of the composite cross-beams passes through a Z-shaped composite stiffener.

16. A method of forming an aircraft empennage, the method comprising:
forming a vertical stabilizer via bonding a composite first understructure to a composite laminate first skin bonded, the composite laminate first skin being:
subject to Mode I, II and III loading; and
comprising a first set of interlaminar fracture toughnesses in Modes I, II, and III; and

forming at least one horizontal stabilizer via bonding a composite second understructure to a composite laminate second skin bonded, the composite laminate second skin being:
subject to Mode I, II and III loading; and
comprising a second set of interlaminar fracture toughnesses in Modes I, II, and III that are lesser in value than the first set of interlaminar fracture toughnesses.

17. The method of claim 16 , wherein the composite laminate first skin further comprises:
a Mode I interlaminar fracture toughness within a range that comprises approximately 4.0 to 6.5 inch-pounds per square inch;
a Mode II interlaminar fracture toughness within a range that comprises approximately 12.0 to 15.5 inch-pounds per square inch; and
a Mode III interlaminar fracture toughness within a range that comprises approximately 16.0 to 18.5 inch-pounds per square inch.

18. The method of claim 16 , wh

Assignees

• Boeing Co

Inventors

• Dan-Jumbo Eugene A

Classifications

• B32B27/20

  using fillers, pigments, thixotroping agents · CPC title

• B64C5/02

  Tailplanes · CPC title

• B64C5/06

  Fins (B64C5/08 takes precedence) · CPC title

• B64C3/182

  Stringers, longerons · CPC title

• B32B2307/546

  Flexural strength; Flexion stiffness · CPC title