Turbine blade having film cooling hole arrangement

Patent metadata

Field	Value
Publication number	US-10036259-B2
Application number	US-201514929692-A
Country	US
Kind code	B2
Filing date	Nov 2, 2015
Priority date	Nov 3, 2014
Publication date	Jul 31, 2018
Grant date	Jul 31, 2018

Abstract

A turbine blade includes a platform that has a platform leading edge and trailing edge joined by two platform circumferential sides. An airfoil extends radially outwardly from the platform to a free tip end. The airfoil includes an airfoil leading edge and trailing edge joined by opposed pressure and suction sides. A root extends radially inwardly from the platform. The platform and the airfoil include film cooling holes arranged in substantial conformance with the Cartesian coordinates set forth in Table 1 herein. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.0 mm).

First claim

What is claimed is:

1. A turbine blade comprising:
a platform that has a platform leading edge and trailing edge joined by two platform circumferential sides;
an airfoil that extends radially outwardly from the platform to a free tip end, the airfoil includes an airfoil leading edge and trailing edge joined by opposed pressure and suction sides; and
a root that extends radially inwardly from the platform, and
the platform and the airfoil include film cooling holes having external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.0 mm).

2. The turbine blade as recited in claim 1 , wherein the turbine blade is a first stage turbine blade.

3. The turbine blade as recited in claim 1 , wherein a portion of the film cooling holes are diffusing and another portion of the film cooling holes are cylindrical.

4. The turbine blade as recited in claim 1 , wherein at least one of the film cooling holes has a hole geometry as set forth in Table 1.

5. The turbine blade as recited in claim 1 , wherein at least one film cooling hole on the airfoil is a diffusing hole and at least one film cooling hole on the platform is a cylindrical hole.

6. The turbine blade as recited in claim 1 , wherein the zero-coordinate is on a surface of the root.

7. The turbine blade as recited in claim 1 , wherein spacing between edges of adjacent cooling holes is at least 0.015 inch (0.38 mm).

8. The turbine blade as recited in claim 1 , wherein the film cooling holes have a diameter of 0.010-0.035 inch (0.25-0.89 mm).

9. A gas turbine engine comprising:
a compressor section;
a combustor fluidly connected to the compressor section;
a turbine section fluidly connected to the combustor, the turbine section includes an array of turbine blades, each turbine blade comprising:
a platform that has a platform leading edge and trailing edge joined by two platform circumferential sides;
an airfoil that extends radially outwardly from the platform to a free tip end, the airfoil includes an airfoil leading edge and trailing edge joined by opposed pressure and suction sides; and
a root that extends radially inwardly from the platform, and
the platform and the airfoil include film cooling holes having external breakout points that are located in substantial conformance with to the Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.0 mm).

10. The gas turbine engine as recited in claim 9 , wherein the turbine blade is a first stage turbine blade.

11. The gas turbine engine as recited in claim 9 , wherein a portion of the film cooling holes are diffusing and another portion of the film cooling holes are cylindrical.

12. The gas turbine engine as recited in claim 9 , wherein at least one of the film cooling holes has a hole geometry as set forth in Table 1.

13. The gas turbine engine as recited in claim 9 , wherein at least one film cooling hole on the airfoil is a diffusing hole and at least one film cooling hole on the platform is a cylindrical hole.

14. The gas turbine engine as recited in claim 9 , wherein the zero-coordinate is on a surface of the root.

15. The gas turbine engine as recited in claim 9 , wherein spacing between edges of adjacent cooling holes is at least 0.015 inch (0.38 mm).

16. The gas turbine engine as recited in claim 9 , wherein the film cooling holes have a diameter of 0.010-0.035 inch (0.25-0.89 mm).

Assignees

• United Technologies Corp

Inventors

• Spangler Brandon W
• Burstynski Laurie L

Classifications

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F01D5/186 Primary

  Film cooling (F01D5/187 takes precedence) · CPC title

• Y02T50/673

  Cross-Sectional Technologies · mapped topic

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title

• F05D2240/81

  Cooled platforms · CPC title