Turbine vane for gas turbine engine

Patent metadata

Field	Value
Publication number	US-10844729-B2
Application number	US-201815946418-A
Country	US
Kind code	B2
Filing date	Apr 5, 2018
Priority date	Apr 5, 2018
Publication date	Nov 24, 2020
Grant date	Nov 24, 2020

Abstract

A turbine vane for a gas turbine engine having a plurality of cooling holes defined therein is provided. The plurality of cooling holes provide fluid communication to a surface of the turbine vane, the plurality of cooling holes including holes noted by the following coordinates: TVA, TVB, TVC, TVD and TVE of Table 1.

First claim

What is claimed is:

1. A turbine vane for a gas turbine engine, comprising:
an inner platform;
an outer platform;
an airfoil extending between the inner platform and the outer platform; and
a plurality of cooling holes defined in the turbine vane, the plurality of cooling holes provide fluid communication to a surface of the turbine vane, the plurality of cooling holes including holes defined by the following coordinates: TVA, TVB, TVC, TVD and TVE of Table 1, wherein for each hole, external breakout center corresponds to an intersection of a central axis of the cooling hole with an outer surface of the vane and wherein the X, Y, Z Cartesian coordinate values of Table 1 have a tolerance of ±0.100 inches of the nominal location with respect to the X, Y and Z axes.

2. The turbine vane of claim 1 , wherein the turbine vane is a first stage turbine vane of a high pressure turbine of a gas turbine engine.

3. The turbine vane of claim 1 , wherein the plurality of cooling holes include holes defined by the all of the coordinates of Table 1.

4. The turbine vane of claim 1 , wherein the plurality of cooling holes include holes defined by the following coordinates of Table 1: TWA, TWB, TWC, TWD, TWE, TWF and TUA, TUB, TUC and TUD.

5. The turbine vane of claim 1 , wherein the turbine vane is a first stage turbine vane of a high pressure turbine of a gas turbine engine.

6. The turbine vane of claim 5 , wherein the plurality of cooling holes include holes defined by the following coordinates of Table 1: TWA, TWB, TWC, TWD, TWE, TWF and TUA, TUB, TUC and TUD.

7. The turbine vane of claim 1 , wherein the plurality of cooling holes are formed by an electrical discharge machining process.

8. The turbine vane of claim 7 , wherein the plurality of cooling holes include holes defined by the following coordinates of Table 1: TWA, TWB, TWC, TWD, TWE, TWF and TUA, TUB, TUC and TUD.

9. The turbine vane of claim 7 , wherein the turbine vane is a first stage turbine vane of a high pressure turbine of a gas turbine engine.

10. The turbine vane of claim 9 , wherein the plurality of cooling holes include holes defined by the all of the coordinates of Table 1.

11. The turbine vane of claim 7 , wherein for each hole, external breakout center corresponds to an intersection of a central axis of the cooling hole with an outer surface of the vane.

12. A turbine vane for a gas turbine engine, comprising:
an inner platform;
an outer platform;
an airfoil extending between the inner platform and the outer platform; and
a plurality of cooling holes defined in the turbine vane, the plurality of cooling holes providing fluid communication to a surface of the turbine vane, the plurality of cooling holes including holes defined by the following coordinates: TWA, TWB, TWC, TWD, TWE, TWF and TUA, TUB, TUC and TUD of Table 1, wherein for each hole, external breakout center corresponds to an intersection of a central axis of the cooling hole with an outer surface of the vane and wherein the X, Y, Z Cartesian coordinate values of Table 1 have a tolerance of ±0.100 inches of the nominal location with respect to the X, Y and Z axes.

13. The turbine vane of claim 12 , wherein the plurality of cooling holes are formed by an electrical discharge machining process.

14. The turbine vane of claim 13 , wherein the turbine vane is a first stage turbine vane of a high pressure turbine of a gas turbine engine.

15. The turbine vane of claim 12 , wherein the turbine vane is a first stage turbine vane of a high pressure turbine of a gas turbine engine.

16. A turbine stator assembly for a gas turbine engine comprising a plurality of vanes, each vane having:
an inner platform;
an outer platform;
an airfoil extending between the inner platform and the outer platform; and
a plurality of cooling holes defined in the turbine vane for providing fluid communication to a surface of each vane, the plurality of cooling holes including holes defined by the following coordinates: TVA, TVB, TVC, TVD and TVE of Table 1, wherein for each hole, external breakout center corresponds to an intersection of a central axis of the cooling hole with an outer surface of the vane and wherein the X, Y, Z Cartesian coordinate values of Table 1 have a tolerance of ±0.100 inches of the nominal location with respect to the X, Y and Z axes.

17. The turbine stator assembly of claim 16 , wherein the plurality of cooling holes include holes defined by the all of the coordinates of Table 1.

18. A method of forming a cooling hole pattern in an exterior surface of a vane used in a high pressure turbine of a gas turbine engine, comprising:
drilling a plurality of cooling holes in the exterior surface of the vane, wherein the plurality of cooling holes provide fluid communication to the exterior surface of the turbine vane, the plurality of cooling holes including holes defined by the following coordinates: TVA, TVB, TVC, TVD and TVE of Table 1, wherein for each hole, external breakout center corresponds to an intersection of a central axis of the cooling hole with an outer surface of the vane and wherein the X, Y, Z Cartesian coordinate values of Table 1 have a tolerance of ±0.100 inches of the nominal location with respect to the X, Y and Z axes.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Arisi Allan N
• Teller Bret M

Classifications

• F05D2250/74

  given by a set or table of xyz-coordinates · CPC title

• F05D2230/11

  by electrochemical methods · CPC title

• F01D5/186 Primary

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

• B23H9/10 Primary

  Working turbine blades or nozzles · CPC title

• Y02T50/60

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