Method and system for interfacing a ceramic matrix composite component to a metallic component

Patent metadata

Field	Value
Publication number	US-10309240-B2
Application number	US-201514808000-A
Country	US
Kind code	B2
Filing date	Jul 24, 2015
Priority date	Jul 24, 2015
Publication date	Jun 4, 2019
Grant date	Jun 4, 2019

Abstract

An airfoil assembly for a gas turbine engine and a method of transferring load from the ceramic matrix composite (CMC) airfoil assembly to a metallic vane assembly support member are provided. The airfoil assembly includes a forward end and an aft end with respect to an axial direction of the gas turbine engine. The airfoil assembly includes a radially outer end component, a radially inner end component, and a hollow airfoil body extending therebetween. The radially outer end component including a radially outwardly-facing end surface having a non-compression load-bearing feature extending radially outwardly and formed integrally with the outer end component, the load-bearing feature configured to mate with a complementary feature formed in a radially inner surface of a first airfoil assembly support structure and selectively positioned orthogonally to a force imparted into the airfoil assembly.

First claim

What is claimed is:

1. An airfoil assembly for a gas turbine engine, said airfoil assembly comprising a ceramic matrix composite (CMC) material, said airfoil assembly comprising a forward end and an aft end with respect to an axial direction of the gas turbine engine, said airfoil assembly comprising:
a radially outer end component comprising a radially outwardly-facing end surface having a forward flange on the forward end, an aft flange on the aft end and a non-compression load-bearing feature extending radially outwardly from said outwardly-facing end surface and formed integrally with said outer end component, said non-compression load-bearing feature located in a portion between the forward flange and the aft flange and configured to mate with a complementary feature formed in a radially inner surface of a first airfoil assembly support structure, said non-compression load-bearing feature selectively positioned orthogonally to a force imparted into said airfoil assembly;
a radially inner end component configured to engage a second airfoil assembly support structure positioned radially inward from said radially inner end component; and
a hollow airfoil body extending there between, said airfoil body configured to receive a strut couplable at a first end to said first airfoil assembly support structure.

2. The assembly of claim 1 , wherein said first airfoil assembly support structure and said second airfoil assembly support structure are composed of a metallic material.

3. The assembly of claim 1 , wherein said radially inner end component comprises a radial retention feature comprising a radial retention pin extending through said radially inner end component and into said second airfoil assembly support structure and configured to maintain a loading between the radially inner end component and the second airfoil assembly support structure such that radially inner end component is clamped to second airfoil assembly support structure.

4. The assembly of claim 1 , wherein said radially inner end component comprises a radially inwardly-facing end surface having a non-compression load-bearing feature extending radially inwardly from said inwardly-facing end surface and formed integrally with said inner end component, said non-compression load-bearing feature of the radially inwardly-facing surface configured to mate with a complementary feature formed in a radially outer surface of a second airfoil assembly support structure.

5. The assembly of claim 4 , wherein said strut is couplable at a second end to said second airfoil assembly support structure.

6. The assembly of claim 1 , wherein said non-compression load-bearing feature comprises a notch formed in a wedge-shaped portion of said outwardly-facing end surface.

7. The assembly of claim 1 , wherein said non-compression load-bearing feature comprises a wedge-shaped portion of said outwardly-facing end surface positioned orthogonally to the axial direction.

8. The assembly of claim 1 , wherein said non-compression load-bearing feature comprises a wedge-shaped portion of said outwardly-facing end surface positioned orthogonally to a circumferential direction approximately orthogonal to the axial direction.

9. The assembly of claim 8 , wherein said wedge-shaped portion engages a pivot member configured to rotate about a radially oriented pin that permits the pivot member to maintain face-to-face contact with said wedge-shaped portion when said airfoil assembly experiences a twisting force.

10. The assembly of claim 1 , wherein said outwardly-facing end surface comprises a plurality of non-compression load-bearing features, each positioned orthogonally to a predetermined direction of a component of a force imparted to said airfoil assembly when said airfoil assembly is in operation within the gas turbine engine.

11. The assembly of claim 1 , wherein said non-compression load-bearing feature comprises an outwardly radially extending tab, said tab configured to engage a complementarily-shaped boss formed in said first airfoil assembly support structure.

12. The assembly of claim 11 , wherein said tab and said boss comprise mutually aligned apertures configured to receive a pin therethrough.

13. The assembly of claim 1 , wherein said non-compression load-bearing feature comprises a hook member comprising a radially outwardly extending ramp portion and an opposing concave portion.

14. The assembly of claim 1 , wherein said outwardly-facing end surface comprises an aperture extending therethrough to an interior of said hollow airfoil body and a pressure-side wedge extending from a pressure-side of said airfoil assembly on said outwardly-facing end surface and terminating at said aperture, said pressure side wedge comprising one or more load pads adjacent said aperture, said one or more load pads configured to receive a complementarily-shaped portion of the first airfoil assembly support structure.

15. A method of transferring load from a ceramic matrix composite (CMC) vane assembly to a metallic vane assembly support member, said method comprising:
providing the CMC vane assembly, the vane assembly including:
a radially outer end component including a radially outwardly facing surface having a forward flange on a forward end, an aft flange on an aft end and one or more radially outwardly extending load transfer features located in a portion between the forward flange and the aft flange;
a radially inner end component; and
an airfoil body extending therebetween;

engaging the radially outer end component to at least one of a plurality of metallic vane assembly support members spaced circumferentially about a gas flow path, the vane assembly support members including one or more load receiving features shaped complementary to the load transfer features, the load transfer feature including a wedge-shaped cross-section.

16. The method of claim 15 , wherein providing the CMC vane assembly comprises providing the CMC vane assembly that includes a second load transfer feature extending radially outwardly from the radially outwardly facing surface of the radially outer end component.

17. A gas turbine engine comprising:
an inner support structure formed of a first metallic material, said inner support structure comprising a strut, said strut comprising a first mating end, a second opposing mating end and a strut body extending radially therebetween;
an outer support structure formed of a second metallic material;
an airfoil assembly comprising a ceramic matrix composite (CMC) material and extending between said inner support structure and said outer support structure, said airfoil assembly comprising:
a radially outer end component comprising a radially outwardly-facing end surface having a forward flange on a forward end, an aft flange on an aft end and a non-compression load-bearing feature extending radially outwardly from said outwardly-facing end surface and formed integrally with said outer end component, said non-compression load-bearing feature located in a portion between the forward flange and the aft flange and configured to mate with a complementary feature formed in a radially inner surface of said outer support structure, said non-compression load-bearing feature selectively positioned orthogonally to a force imparted into said radially outwardly-facing end surface;
a radially inner end component; and
a hollow airfoil body extending therebetween, said airfoil body configured to receive a strut couplable at a first end to said outer support structure.

18. The gas turbine engine of claim 17 , wherein said

Assignees

• Gen Electric

Inventors

• Heitman Bryce Loring
• Senile Darrell Glenn
• Tuertscher Michael Ray
• Phelps Greg Scott
• Feie Brian Gregg
• Murphy Steven James

Classifications

• F01D9/042 Primary

  fixing blades to stators (fixing stator-rings in the casing or to each other F01D25/246) · CPC title

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title

• F01D5/18

  Hollow blades, {i.e. blades with cooling or heating channels or cavities (structure of hollow blades in general F01D5/147)}; Heating, heat-insulating or cooling means on blades · CPC title

• F01D9/041 Primary

  using blades (F01D5/148 takes precedence) · CPC title

• F05D2220/32

  in gas turbines · CPC title