Closed-angle composite airfoil spar and method of fabricating the same

Patent metadata

Field	Value
Publication number	US-10647406-B2
Application number	US-201715610922-A
Country	US
Kind code	B2
Filing date	Jun 1, 2017
Priority date	Jun 1, 2017
Publication date	May 12, 2020
Grant date	May 12, 2020

Abstract

A closed-angle composite airfoil spar is provided for an airfoil assembly. The composite airfoil spar comprises a web portion having a thickness. The composite airfoil spar also comprises a flange portion having a thickness. The flange portion extends transverse to the web portion. The composite airfoil spar further comprises a radius portion interconnecting the web and flange portions. The radius portion is an acceptable amount thinner relative to thickness of the web portion and thickness of the flange portion based on design requirements and material properties.

First claim

What is claimed is:

1. A closed-angle composite airfoil spar for an airfoil assembly, the composite airfoil spar comprising:
a web portion having a web portion thickness;
an elongated flange portion having an elongated flange portion thickness and extending transverse to the web portion and forming an acute angle between the web portion and the elongated flange portion, wherein the elongated flange portion serves as a splice between a leading edge and a skin panel of the airfoil assembly, and wherein the web portion thickness is substantially equal to the elongated flange portion thickness; and
a radius portion interconnecting the web portion and the elongated flange portion, wherein the radius portion has a radius portion thickness, and wherein the radius portion thickness is no more than 10 percent thinner than the web portion thickness and no more than 10 percent thinner than the elongated flange portion thickness.

2. The closed-angle composite airfoil spar of claim 1 wherein the radius portion thickness is no more than 5 percent thinner than the web portion thickness and no more than 5 percent thinner than the elongated flange portion thickness.

3. The closed-angle composite airfoil spar of claim 1 wherein the acute angle is at most 80 degrees.

4. The closed-angle composite airfoil spar of claim 1 having a cross-sectional shape resembling the letter “C”.

5. The closed-angle composite airfoil spar of claim 1 wherein the web portion thickness is 0.220+/−0.005 inches (5.6+/−0.1 millimeters), the elongated flange portion thickness is 0.220+/−0.005 inches (5.6+/−0.1 millimeters), and the radius portion thickness is 0.210+/−0.005 inches (5.3+/−0.1 millimeters).

6. The closed-angle composite airfoil spar of claim 1 wherein the radius portion thickness is no more than 5 percent thinner than the web portion thickness and no more than 5 percent thinner than the elongated flange portion thickness.

7. The closed-angle composite airfoil spar of claim 1 wherein the flange portion comprises a joggle that is elongated along a longitudinal length of the flange portion.

8. The closed-angle composite airfoil spar of claim 1 wherein the splice is between a leading edge and an upper skin panel of the airfoil assembly.

9. The closed-angle composite airfoil spar of claim 8 wherein the leading edge meets the upper skin panel at a junction, and wherein the web portion is located aft of the junction.

10. The closed-angle composite airfoil spar of claim 1 wherein the material of the web portion, the flange portion, and the radius portion comprises a fiber reinforced material.

11. The closed-angle composite airfoil spar of claim 10 wherein the fiber reinforced material comprises an epoxy material.

12. The closed-angle composite airfoil spar of claim 11 wherein the epoxy material comprises epoxy resin that gels at a temperature between 240° F. and 260° F.

13. The closed-angle composite airfoil spar of claim 11 wherein the fiber reinforced material further comprises carbon fibers.

14. An airfoil assembly comprising:
a leading edge;
a skin panel abutting the leading edge; and
a closed-angle spar comprising:
a web portion having a web portion thickness;
an elongated flange portion having an elongated flange portion thickness that is substantially equal to the web portion thickness, the elongated flange portion extending transverse to the web portion and forming an acute angle between the web portion and the elongated flange portion, wherein the elongated flange portion serves as a splice between the leading edge and the skin panel; and
a radius portion interconnecting the web portion and the elongated flange portion, the radius portion having a radius portion thickness that is no more than 10 percent thinner than the web portion thickness and no more than 10 percent thinner than the elongated flange portion thickness.

15. The airfoil assembly of claim 14 wherein the acute angle faces forward of the airfoil assembly after the leading edge and the skin panel are spliced together.

16. The airfoil assembly of claim 14 wherein the web portion is located aft of the splice point after the leading edge and the skin panel are spliced together along the elongated flange portion.

17. The airfoil assembly of claim 14 wherein the closed-angle spar is a forward-facing spar for a front portion of the airfoil assembly.

18. The airfoil assembly of claim 14 wherein the acute angle is at most 80 degrees.

19. The airfoil assembly of claim 14 wherein the web portion thickness is 0.220+/−0.005 inches (5.6+/−0.1 millimeters), the elongated flange portion thickness is 0.220+/−0.005 inches (5.6+/−0.1 millimeters), and the radius portion thickness is 0.210+/−0.005 inches (5.3+/−0.1 millimeters).

20. The airfoil assembly of claim 14 wherein the material of the web portion, the flange portion, and the radius portion comprises a fiber reinforced material.

21. The airfoil assembly of claim 20 wherein the fiber reinforced material comprises an epoxy material and carbon fibers.

22. The airfoil assembly of claim 21 wherein the epoxy material comprises epoxy resin that gels at a temperature between 240° F. and 260° F.

23. A method of fabricating the closed-angle composite airfoil spar of claim 1 comprising:
applying layers of a fiber reinforced material to a tooling piece to build up layers of the fiber reinforced material on the tooling piece, the tooling piece comprising a plurality of temporarily joined mandrel pieces;
drawing a vacuum about the tooling piece including the layers of fiber reinforced material on the tooling piece for at least a first predetermined process time;
heating the tooling piece including the layers of the fiber reinforced material on the tooling piece for at least a second predetermined process time; and
after elapse of the first predetermined process time, releasing the vacuum drawn about the tooling piece including the layers of fiber reinforced material on the tooling piece to produce the closed-angle composite airfoil spar having a closed angle no more than 80 degrees.

24. The method of claim 23 wherein the heating the tooling piece including the layers of the fiber reinforced material on the tooling piece for at least the second predetermined process time includes heating the tooling piece including the layers of the fiber reinforced material on the tooling piece using an out-of-autoclave process.

25. The method of claim 23 wherein the drawing the vacuum about the tooling piece including the layers of fiber reinforced material on the tooling piece for at least the first predetermined process time includes drawing a vacuum about the tooling piece including the layers of fiber reinforced material on the tooling piece for a minimum time specified by a manufacturer of the fiber reinforced material.

26. The method of claim 23 wherein the applying the layers of the fiber reinforced material to the tooling piece to build up the layers of the fiber reinforced material on the tooling piece includes applying the layers of the fiber reinforced material to the tooling piece to build up a minimum number of layers of the fiber reinforced material on the tooling piece.

27. The method of claim 23 wherein the applying the layers of the fiber reinforced material to the tooling piece comprises selecting a fiber reinforced material comprising an epoxy material.

28. The metho

Assignees

• Boeing Co

Inventors

• Deck Eric E
• Rhodes Charles Alan
• Berner Jeffry K

Classifications

• B29L2031/3085

  Wings · CPC title

• B64C3/20

  Integral or sandwich constructions · CPC title

• B64C3/185 Primary

  Spars · CPC title

• B64C3/26

  Construction, shape, or attachment of separate skins, e.g. panels · CPC title

• F04D29/668

  damping or preventing mechanical vibrations · CPC title