Cooling schemes for airfoils for gas turbine engines

Patent metadata

Field	Value
Publication number	US-11629602-B2
Application number	US-202217841839-A
Country	US
Kind code	B2
Filing date	Jun 16, 2022
Priority date	Jun 17, 2021
Publication date	Apr 18, 2023
Grant date	Apr 18, 2023

Abstract

Airfoil for gas turbine engines are described. The airfoils have internal walls and internal cavities. A leading edge cavity is defined within the airfoil body along a leading edge and a leading edge feed cavity is arranged aft of the leading edge cavity. A bent leading edge rib is arranged between the leading edge cavity and the leading edge feed cavity. A main body cavity is arranged aft of the leading edge feed cavity and defined at least in part by two interior ribs that define a part of the leading edge feed cavity. The main body cavity is fluidly connected to the leading edge feed cavity by an interior fluid connection through the intersection of the two interior ribs. A shield cavity is arranged to thermally shield the leading edge feed cavity from heat pickup along the suction side of the airfoil body.

First claim

What is claimed is:

1. An airfoil for a gas turbine engine, the airfoil comprising:
an airfoil body having a leading edge, a trailing edge, a pressure side exterior wall extending between the leading edge and the trailing edge and defining a pressure side, and a suction side exterior wall extending between the leading edge and the trailing edge and defining a suction side, wherein a plurality of cooling passages are formed within the airfoil;
a plurality of interior ribs defining one or more skin core cavities and one or more internal air passages, wherein the plurality of interior ribs form a flexible structure that is configured to reduce thermal fight and stresses within material of the airfoil body, wherein each of the plurality of interior ribs extends between the pressure side exterior wall and the suction side exterior wall;
a leading edge cavity defined within the airfoil body and defined along the leading edge to provide cooling to the leading edge of the airfoil;
a leading edge feed cavity arranged aft of the leading edge cavity and defined in part by two interior ribs that extend from and cross between the pressure side and the suction side and intersect each other at a point between the pressure side and the suction side;
a leading edge rib arranged between the leading edge cavity and the leading edge feed cavity, wherein the leading edge rib is a bent rib having an apex, a pressure side leg extending from the apex to the pressure side of the airfoil body, a suction side leg extending from the apex to the suction side of the airfoil body, and at least one impingement hole fluidly connecting the leading edge feed cavity to the leading edge cavity;
a main body cavity arranged aft of the leading edge feed cavity and defined at least in part by the two interior ribs that define a part of the leading edge feed cavity, wherein the main body cavity is fluidly connected to the leading edge feed cavity by an interior fluid connection through the intersection of the two interior ribs; and
a feed cavity-shield cavity arranged along the suction side of the airfoil body, forward of the main body cavity, aft of the leading edge cavity, and configured to thermally shield the leading edge feed cavity from heat pickup along the suction side of the airfoil body.

2. The airfoil of claim 1 , further comprising a plurality of suction side skin core cavities arranged along the suction side of the airfoil body aft of the leading edge feed cavity and defined by the plurality of interior ribs.

3. The airfoil of claim 1 , further comprising a plurality of pressure side skin core cavities arranged along the pressure side of the airfoil body aft of the leading edge feed cavity and defined by the plurality of interior ribs.

4. The airfoil of claim 1 , wherein the interior fluid connection is a slot extending a full span of the airfoil body in a radial direction.

5. The airfoil of claim 1 , wherein the interior fluid connection is a slot extending less than a full span of the airfoil body in a radial direction.

6. The airfoil of claim 5 , wherein the slot extends for 50% of a radial span of the airfoil body.

7. The airfoil of claim 5 , wherein a solid rib extends from a root of the airfoil to the slot and the slot extends from the solid rib to the end of the full span of the airfoil.

8. The airfoil of claim 1 , wherein the interior fluid connection comprises at least one crossover aperture formed in a rib between the leading edge feed cavity and the main body cavity.

9. The airfoil of claim 1 , further comprising at least one pedestal spanning the fluid connection such that a mechanical connection between the pressure side and the suction side is provided through the at least one pedestal.

10. The airfoil of claim 1 , wherein the one or more skin core cavities are triangular in shape in cross-section.

11. The airfoil of claim 1 , wherein the apex is arranged coincident with a camber line of the airfoil body and the pressure side leg extends toward the trailing edge at an angle of between 30° and 75° relative to a tangent line of the camber line at the apex.

12. The airfoil of claim 1 , wherein the apex is arranged coincident with a camber line of the airfoil body and the suction side leg extends toward the trailing edge at an angle of between 30° and 75° relative to a tangent line of the camber line at the apex.

13. The airfoil of claim 11 , wherein each of the preside side leg and the suction side leg extend toward the trailing edge of the airfoil body with an angle of between 60° and 150° between the pressure side leg and the suction side leg.

14. The airfoil of claim 1 , wherein a circumferential direction dimension of the internal air passages and a circumferential direction dimension of the skin core cavities are sized such that adjacent internal air passages overlap in a circumferential direction with at least a portion of the skin core cavities.

15. A gas turbine engine comprising:
a turbine section;
a compressor section; and
an airfoil located in at least one of the turbine section and the compressor section, the airfoil comprising:
an airfoil body having a leading edge, a trailing edge, a pressure side exterior wall extending between the leading edge and the trailing edge and defining a pressure side, and a suction side exterior wall extending between the leading edge and the trailing edge and defining a suction side, wherein a plurality of cooling passages are formed within the airfoil;
a plurality of interior ribs defining one or more skin core cavities and one or more internal air passages, wherein the plurality of interior ribs form a flexible structure that is configured to reduce thermal fight and stresses within material of the airfoil body, wherein each of the plurality of interior ribs extends between the pressure side exterior wall and the suction side exterior wall;
a leading edge cavity defined within the airfoil body and defined along the leading edge to provide cooling to the leading edge of the airfoil;
a leading edge feed cavity arranged aft of the leading edge cavity and defined in part by two interior ribs that extend from and cross between the pressure side and the suction side and intersect each other at a point between the pressure side and the suction side;
a leading edge rib arranged between the leading edge cavity and the leading edge feed cavity, wherein the leading edge rib is a bent rib having an apex, a pressure side leg extending from the apex to the pressure side of the airfoil body, a suction side leg extending from the apex to the suction side of the airfoil body, and at least one impingement hole fluidly connecting the leading edge feed cavity to the leading edge cavity;
a main body cavity arranged aft of the leading edge feed cavity and defined at least in part by the two interior ribs that define a part of the leading edge feed cavity, wherein the main body cavity is fluidly connected to the leading edge feed cavity by an interior fluid connection through the intersection of the two interior ribs; and
a feed cavity-shield cavity arranged along the suction side of the airfoil body, forward of the main body cavity, aft of the leading edge cavity, and configured to thermally shield the leading edge feed cavity from heat pickup along the suction side of the airfoil body.

16. The gas turbine engine of claim 15 , wherein the interior fluid connection is one of a slot extending a full span of the airfoil body in a radial direction and a slot extending less than a full span of the airfoil body in a radial direction.

17. The gas turbine engine of claim

Assignees

• Raytheon Tech Corp

Inventors

• Spangler Brandon W
• Pack David R

Classifications

• F01D5/187 Primary

  Convection cooling · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F05D2240/303

  related to the leading edge of a rotor blade · CPC title

• F05D2240/121

  related to the leading edge of a stator vane · CPC title

• F05D2260/202

  by film cooling · CPC title