Method for reducing cooled turbine element stress and element made thereby

• US 6,514,037 B1
• Filed: 09/26/2001
• Issued: 02/04/2003
• Est. Priority Date: 09/26/2001
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Danowski, Michael Joseph, Brainch, Gulcharan Singh
• Primary Examiner(s): VERDIER, CHRISTOPHER M
• Application Number:
  US09/964,040
• Time in Patent Office: 496 Days
• Field of Search: 415/1, 415/115, 415/116, 416/1, 416/96.R, 416/96.A, 416/97.R
• US Class Current: 415/115
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2240/81 Cooled platforms
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/221 Improvement of heat transfer
• CPC Class Codes: F05D 2270/114 by limiting mechanical stresses

Abstract

A cooled turbine element including an airfoil and a flowpath boundary member extending laterally from either an inboard end or an outboard end of the airfoil. The member has a flowpath face and an outside face which is cooler than said flowpath face creating a tendency for the member to deflect in a direction away from the flowpath face and causing a thermally induced tensile radial stress in a region of the trailing edge of the airfoil. The element has an interior cooling passage and at least one cooling hole extending from the interior cooling passage to an opening located in an area upstream from the stressed region of the trailing edge to cool the area so the airfoil thermally deflects to a shape corresponding to that of the boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.

Claims

1. A method of lowering a thermal stress at a trailing edge of an airfoil of a cooled turbine blade adjacent a platform of the blade, said method comprising the step of forming at least one cooling hole positioned upstream from the trailing edge of the airfoil and extending from an interior cooling air passage to an exterior surface of the airfoil for delivering cooling air to the exterior surface to cool the airfoil in an area of the exterior surface upstream from the trailing edge so that a thermal deflection of the airfoil more closely corresponds to a thermal deflection of the platform thereby lowering thermally induced stresses in the airfoil at the trailing edge thereof.

2. A method as set forth in claim 1 wherein said at least one cooling hole is formed on a pressure side of the airfoil so that the thermal deflection of the airfoil more closely corresponds to the thermal deflection of the platform to lower thermally induced bending stresses in the airfoil at the trailing edge thereof.

4. An element as set forth in claim 1 wherein the element is a cooled turbine blade and the lateral boundary member is a platform thereof positioned at the inboard end of the airfoil.

5. An element as set forth in claim 1 wherein the cooling hole extends to said pressure side of the airfoil.

6. An element as set forth in claim 5 wherein the cooling hole extends at an angle of between about twenty degrees and about forty degrees with respect to said pressure side of the airfoil.

7. An element as set forth in claim 1 wherein the position to which the cooling hole extends is located on the airfoil between about 65 percent chord and about 85 percent chord.

8. An element as set forth in claim 7 wherein the position to which the cooling hole extends is located on the airfoil between about seventy percent chord and about 83 percent chord.

9. An element as set forth in claim 1 wherein the position to which the cooling hole extends is located on the airfoil between about zero percent span and about ten percent span.

10. An element as set forth in claim 9 wherein the position to which the cooling hole extends is located on the airfoil between about four percent span and about six percent span.

11. An element as set forth in claim 1 wherein the cooling hole extends radially outward at an angle of between about zero degrees and about ninety degrees with respect to an axial direction of the engine.

12. An element as set forth in claim 1 wherein the cooling hole diverges from the interior cooling passage to the position.

13. An element as set forth in claim 12 wherein the cooling hole diverges at an angle of between about zero degrees and about twenty degrees.

14. An element as set forth in claim 1 wherein the element has four cooling holes extending from the interior cooling passage to positions located in the area to cool said area to a temperature below that of the trailing edge so that the airfoil thermally deflects during engine operation to a shape corresponding to that of the flowpath boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.

3. A cooled turbine element for use in a flowpath of a gas turbine engine comprising:

an airfoil having a pressure side and a suction side opposite said pressure side, said pressure side and said suction side extending axially between a leading edge and a trailing edge opposite said leading edge and radially between an inboard end and an outboard end opposite said inboard end;

a flowpath boundary member extending laterally from at least one of said inboard end and said outboard end, said boundary member having a flowpath face and an outside face opposite the flowpath face, said outside face running cooler than said flowpath face during engine operation thereby creating a tendency for the member to deflect in a direction away from the flowpath face and causing a thermally induced tensile radial stress in a region of the trailing edge of the airfoil;

an interior cooling passage extending through the airfoil from a cooling air source for transporting cooling air through the airfoil; and

at least one cooling hole extending from the interior cooling passage to an opening located on one of said suction side and said pressure side in an area upstream from the stressed region of said trailing edge to cool said area to a temperature below that of the trailing edge so that the airfoil thermally deflects during engine operation to a shape corresponding to that of the flowpath boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.