Hybridized airfoil for a gas turbine engine

Patent metadata

Field	Value
Publication number	US-11248477-B2
Application number	US-201916530917-A
Country	US
Kind code	B2
Filing date	Aug 2, 2019
Priority date	Aug 2, 2019
Publication date	Feb 15, 2022
Grant date	Feb 15, 2022

Abstract

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, a first portion welded to a second portion along an interface such that at least the first portion establishes an airfoil section and the second portion establishes a root section mountable to a rotatable hub. The airfoil section includes an airfoil body extending between leading and trailing edges in a chordwise direction, extending between pressure and suction sides separated in a thickness direction, and extending from the root section in a spanwise direction to a tip portion. A recessed region extends inwardly from at least one of the pressure and suction sides. The airfoil body includes at least one rib bounding a respective pocket within a perimeter of the recessed region. A cover skin is welded to the airfoil body along the at least one rib to enclose the recessed region. A method of forming a gas turbine engine component is also disclosed.

First claim

What is claimed is:

1. An airfoil for a gas turbine engine comprising:
a first portion joined to a second portion along an interface such that at least the first portion establishes an airfoil section and the second portion establishes a root section mountable to a rotatable hub, and wherein:
the airfoil section includes an airfoil body extending between leading and trailing edges in a chordwise direction, extending between pressure and suction sides separated in a thickness direction, and extending from the root section in a spanwise direction to a tip portion;
the tip portion defines a stagger angle relative to the root section, and the stagger angle is greater than or equal to 5 degrees, absolute, prior to joining the first portion to the second portion along the interface;
a recessed region extends inwardly from at least one of the pressure and suction sides; and
the airfoil body includes a plurality of ribs bounding respective pockets within a perimeter of the recessed region, and the plurality of ribs are spaced apart from each other and from the perimeter of the recessed region; and

a cover skin coupled to the airfoil body along the perimeter of the recessed region and along the plurality of ribs to enclose the recessed region.

2. The airfoil as recited in claim 1 , wherein the first portion is linear friction welded to the second portion along the interface.

3. The airfoil as recited in claim 1 , wherein the recessed region is dimensioned to extend across the interface.

4. The airfoil as recited in claim 3 , wherein at least one of the plurality of ribs is dimensioned to extend across the interface.

5. The airfoil as recited in claim 1 , wherein the first portion and the second portion are metallic.

6. The airfoil as recited in claim 5 , wherein the first portion has a first microstructure, and the second portion has a second, different microstructure.

7. The airfoil as recited in claim 1 , wherein the first portion is joined to the second portion by welding along the interface, and wherein the stagger angle is greater than or equal to 15 degrees, absolute, prior to welding the first portion to the second portion.

8. The airfoil as recited in claim 1 , wherein the airfoil is a fan blade.

9. The airfoil as recited in claim 8 , wherein:
the first portion and the second portion are formed from wrought metal comprising titanium such that the first portion has a first microstructure and such that the second portion has a second, different microstructure; and
the stagger angle is greater than or equal to 10 degrees, absolute, and is less than or equal to 30 degrees, absolute, prior to linear friction welding the first portion to the second portion along the interface.

10. A gas turbine engine comprising:
a fan section including a fan;
a compressor section;
a turbine section that drives the compressor section and the fan; and
a plurality of airfoils, wherein each of the plurality of airfoils includes a first portion joined to a second portion along an interface such that at least the first portion establishes an airfoil section including an airfoil body and the second portion establishes a root section;
wherein the first portion is dimensions such that the airfoil body is twisted along a spanwise axis to define a stagger angle prior to joining the first portion to the second portion along the interface;
wherein the airfoil body includes a plurality of ribs bounding respective pockets within a perimeter of a recessed region, and the plurality of ribs are spaced apart from each other and from the perimeter of the recessed region; and
a cover skin coupled to the airfoil body along the perimeter of the recessed region and along the plurality of ribs to enclose the recessed region along one of a pressure side and a suction side of the airfoil section.

11. The gas turbine engine as recited in claim 10 , wherein the fan includes a rotatable hub, and the root section of each of the plurality of airfoils is mounted to the hub.

12. The gas turbine engine as recited in claim 11 , wherein the recessed region is dimensioned to extend across the interface.

13. The gas turbine engine as recited in claim 12 , wherein at least one of the plurality of ribs and at least one of the respective pockets are dimensioned to extend across the interface.

14. The gas turbine engine as recited in claim 11 , wherein the stagger angle is greater than or equal to 15 degrees, absolute, prior to joining the first portion to the second portion along the interface.

15. A method of forming an airfoil for a gas turbine engine comprising:
forming a first portion to establish an airfoil section including an airfoil body, the airfoil body extending along a spanwise axis to a tip portion;
forming a second portion to establish a root section;
joining the first portion to the second portion along an interface, subsequent to twisting the airfoil body along the spanwise axis, such that the tip portion defines a stagger angle relative to the root section;
forming a recessed region along a sidewall of the airfoil body;
forming a plurality of ribs in the airfoil body that bound respective pockets within a perimeter of the recessed region, the plurality of ribs spaced apart from each other and from the perimeter of the recessed region; and
coupling a cover skin to the airfoil body along the perimeter of the recessed region and along the plurality of ribs to enclose the recessed region.

16. The method as recited in claim 15 , wherein the step of joining the first portion to the second portion includes fusion welding the first and second portions along the interface.

17. The method as recited in claim 16 , wherein:
the step of twisting the airfoil body includes hot forming the first portion such that the stagger angle is greater than or equal to 5 degrees, absolute;
the step of forming the first portion includes establishing a first microstructure from wrought metal; and
the step of forming the second portion includes establishing a second, different microstructure from wrought metal.

18. The method as recited in claim 15 , wherein
the step of joining includes welding the first portion to the second portion along the interface, and the step of forming the recessed region occurs subsequent to the step of welding the first and second portions.

19. The method as recited in claim 18 , wherein the step of forming the recessed region includes dimensioning the recessed region to extend across the interface.

20. The method as recited in claim 15 , wherein the airfoil is a fan blade.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Bales Daniel A
• Novikov Dmitri
• Dolan Robert C
• Malmborg Eric W
• Gordon Steven T

Classifications

• F04D29/388

  characterised by construction · CPC title

• F01D5/147 Primary

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title

• B23K20/1205

  using translation movement · CPC title

• F05D2240/301

  Cross-sectional characteristics · CPC title

• B23K2101/001

  Turbines · CPC title