Bonded Composite Aircraft Wing

• US 20140061385A1
• Filed: 08/28/2012
• Published: 03/06/2014
• Est. Priority Date: 08/28/2012
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Dan-Jumbo, Eugene
• Granted Patent: US 8,851,422 B2
• Time in Patent Office: Days
• US Class Current: 244/123.1
• CPC Class Codes: B29C 70/88 characterised primarily by possessing specific properties, e.g. electrically conductive or locally reinforced
• CPC Class Codes: B64C 3/18 Spars; Ribs; Stringers
• CPC Class Codes: B64C 3/185 Spars
• CPC Class Codes: B64C 3/187 Ribs
• CPC Class Codes: B64C 3/20 Integral or sandwich constructions
• CPC Class Codes: B64C 3/26 Construction, shape, or attachment of separate skins, e.g. panels
• CPC Class Codes: Y02T 50/40 Weight reduction

Abstract

A unitized composite aircraft wing includes upper and lower composite laminate wing skins that have differing stiffnesses and are bonded to a composite grid structure. The grid structure includes grid spars extending in a span-wise direction, and cross beams extending in a chord-wise direction of the wing. The grid spars include caps bonded to the upper and lower wing skins.

Claims

1. An aircraft wing, comprising:

upper and lower composite wing skins respectively having differing stiffnesses; and

a wing-grid structure located between and attached to each of the upper and lower wing skins.

2. The aircraft wing of claim 1, wherein:

the wing-grid structure includes a plurality of intersecting composite support members bonded together.

3. The aircraft wing of claim 2, wherein:

the composite support members include wing-grid spars extending in a span-wise direction of the wing and bonded to each of the upper and lower composite wing skins, and

the upper and lower composite wing skins and the wing-grid structure form a unitized bonded wing structure.

4. The aircraft wing of claim 1, wherein:

the upper composite wing skin has a stiffness less than the stiffness of the lower composite wing skin.

5. The aircraft wing of claim 1, wherein:

each of the upper and lower wing skins is orthotropic and has an interlaminar fracture toughness, and

wherein the interlaminar fracture toughness of the upper composite wing skin is less than the interlaminar fracture toughness of the lower composite wing skin.

6. The aircraft wing of claim 5, wherein:

the upper composite wing skin has an interlaminar fracture toughness in Mode I of between approximately 3.0 in-lbs/in² and approximately 5.0 in-lbs/in², and

the lower composite wing skin has an interlaminar fracture toughness in Mode I between approximately 4.5 in-lbs/in² and approximately 6.5 in-lbs/in².

7. The aircraft wing of claim 5 wherein the upper and lower composite wing skins having at least one differing characteristic selected from the group comprising:

fiber orientation,

fiber material,

fiber diameter,

fiber density,

resin material,

number of plies having a fiber orientation of 0 degrees.

8. The aircraft wing of claim 2, wherein the composite support members include:

a plurality of composite wing-grid spars extending in a span-wise direction of the wing, and

a plurality of composite grid cross beams bonded to the wing-grid spars and extending in a chord-wise direction of the wing.

9. The aircraft wing of claim 8, wherein:

each of the wing-grid spars includes a pair of spar caps and a spar web connected between the spar caps, and

the grid cross beams include cross beam webs coupled with the spar webs.

10. The aircraft wing of claim 8 wherein the composite cross beams include:

a first set of substantially flat straps sandwiched between the wing-grid spars and the upper composite wing skin, and

a second set of substantially flat straps sandwiched between the wing-grid spars and the lower composite wing skin.

11. The aircraft wing of claim 1, wherein:

the grid structure includes a plurality of wing-grid spars extending in a span-wise direction of the wing,

each of the wing-grid spars including an upper cap bonded to the upper composite wing skin and a lower cap bonded to the lower composite wing skin.

12. A unitized composite aircraft wing, comprising:

an upper composite wing skin including multiple laminate plies of fiber reinforced resin and having a first interlaminar fracture toughness;

a lower composite wing skin including a plurality of laminated plies of fiber reinforced resin and having a second interlaminar fracture toughness that is greater in magnitude than the first interlaminar fracture toughness; and

a composite grid structure sandwiched between and bonded to each of the upper and lower composite wing skins, the composite grid structure including a plurality of composite wing-grid spars extending in a span-wise direction of the wing, and a plurality of composite grid cross beams cocured with the wing-grid spars and extending in a chord-wise direction of the wing, wherein the composite wing-grid spars are bonded to the upper and lower composite wing skins to form a unitized wing structure.

13. The unitized composite aircraft wing of claim 12, wherein each of the wing-grid spars includes an I-beam having an upper cap bonded to the upper composite wing skin, a lower cap bonded to the lower composite wing skin and a web connecting the upper and lower caps.

14. The unitized composite aircraft wing of claim 12, wherein:

the upper composite wing skin has an interlaminar fracture toughness in Mode I of between approximately 3.0 in-lbs/in² and approximately 5.0 in-lbs/in², in Mode II of between approximately 4.5 in-lbs/in² and 7.0 in-lbs/in², and in Mode III of between approximately 7.5 in-lbs/in² and 8.5 in-lbs/in², and

the lower composite wing skin has an interlaminar fracture toughness in Mode I between approximately 4.5 in-lbs/in² and approximately 6.5 in-lbs/in², in Mode II of between approximately 5.5 in-lbs/in² and 8.0 in-lbs/in², and in Mode III of between approximately 8.5 in-lbs/in² and 12.0 in-lbs/in²,

15. The unitized composite aircraft wing of claim 12, wherein:

each of the wing-grid spars has a grid spar web,

each of the grid cross beams has a cross beam web, and the grid spar webs are connected to the cross beam webs.

16. The unitized composite aircraft wing of claim 12, wherein each of the wing-grid spars includes:

a central web,

upper and lower caps respectively on opposite ends of the central web,

side webs spaced from and extending generally parallel to the central web,

a center stiffener connected to the central web and extending between the side webs.

17. The unitized composite aircraft wing of claim 16, wherein the center stiffener extends between and structurally connects the grid spars.

18. The unitized composite aircraft wing of claim 12, wherein the grid cross beams include first and second sets of substantially flat composite straps respectively sandwiched between the wing-grid spars and the upper and lower composite wing skins.

19. A method of making an aircraft wing structure, comprising:

forming upper and lower composite laminate wing skins respectively having differing stiffnesses;

forming a composite inner wing-grid structure; and

bonding the wing-grid structure to the upper and lower composite wing laminate skins.

20. The method of claim 19, wherein forming the wing-grid structure includes:

forming a plurality of composite wing-grid spars,

forming a plurality of composite grid cross beams, and

bonding the composite wing-grid spars to the composite grid cross beams.

21. The method of claim 19, wherein forming the upper and lower composite laminate wing skins includes selecting at least one material characteristic that results in the upper composite laminate wing skin having an interlaminar fracture toughness that is less than the interlaminar fracture toughness of the lower composite laminate wing skin.

22. The method of claim 21, wherein the material characteristic is selected from the group consisting of:

fiber orientation,

fiber material,

fiber diameter,

fiber density,

resin material,

number of plies having a fiber orientation of 0 degrees.

23. The method of claim 19, wherein:

forming the upper composite laminate wing skin includes selecting an interlaminar fracture toughness in Mode I for the upper composite laminate wing skin that is between approximately 3.0 in-lbs/in² and approximately 5.0 in-lbs/in², and

forming the lower composite laminate wing skin includes selecting an interlaminar fracture toughness in Mode I for the lower composite laminate wing skin that is between approximately 4.5 in-lbs/in² and approximately 6.5 in-lbs/in².

24. The method of claim 19, wherein:

forming the grid structure includes forming wing-grid spars each having a web and upper and lower caps, and

bonding the wing-grid structure to the upper and lower composite laminate wing skins includes bonding the upper and lower caps of each of the wing-grid spars respectively to the upper and lower composite laminate wing skins.

25. The method of claim 19, wherein:

forming upper and lower composite laminate wing skins includes cocuring each of the composite laminate wing skins,

forming the composite inner wing-grid structure includes cocuring a grid of composite structural members, and

bonding the wing-grid structure to the upper and lower composite laminate wing skins includes bonding the cocured upper and lower composite laminate wing skins to the cocured grid of composite structural members.