Gas turbine engine turbine blade tip cooling

Patent metadata

Field	Value
Publication number	US-2016265366-A1
Application number	US-201415032736-A
Country	US
Kind code	A1
Filing date	Oct 10, 2014
Priority date	Nov 11, 2013
Publication date	Sep 15, 2016
Grant date	—

Abstract

A gas turbine engine component includes a structure having a surface configured to be exposed to a hot working fluid. The surface includes a recessed pocket that is circumscribed by an overhang. At least one cooling groove is provided by the overhang.

First claim

What is claimed is:

1 . A gas turbine engine component comprising:
a structure having a surface configured to be exposed to a hot working fluid, the surface includes a recessed pocket that is circumscribed by an overhang, and at least one cooling groove provided by the overhang.

2 . The cooling channel in claim 1 which exits through a continuous channel into the recessed pocket.

3 . The cooling channel in claim 1 which exits through discontinuous channels into the recessed pocket.

4 . The component according to claim 1 , comprising at least one discrete hole in fluid communication with the groove and configured to provide a cooling fluid to the pocket.

5 . The component according to claim 1 , wherein the structure is an instrumentation probe.

6 . The component according to claim 1 , wherein the structure is an airfoil.

7 . The component according to claim 6 , wherein the airfoil includes a cast first portion, and a second portion is secured to the first portion, the second portion providing the overhang.

8 . The component according to claim 7 , wherein the second portion is additively manufactured.

9 . The component according to claim 1 , wherein the overhang circumscribes the pocket.

10 . The component according to claim 9 , wherein the overhang includes a lip that provides an interior perimeter of the pocket, the groove is provided between the overhang and the end wall, the groove bounded by the lip.

11 . The component according to claim 10 , wherein the overhang substantially encloses the groove and provides an exit that fluidly interconnects the groove with the pocket.

12 . The component according to claim 11 , wherein the exit is provided radially between the lip and the end wall.

13 . The component according to claim 9 , wherein the pocket is teardrop-shaped.

14 . The component according to claim 1 , wherein the overhang and an adjacent wall encloses the groove.

15 . A method of manufacturing a turbine blade airfoil, comprising the step of:
(a) forming a structure having a surface configured to be exposed to a hot working fluid (b) forming a surface comprising a recessed pocket; (c) forming an overhang that circumscribes the recessed pocket which includes at least one cooling groove provided by the overhang; and (d) using an additive manufacturing process to create a negative for casting of features for at least one of the steps (a)-(c).

16 . The method according to claim 15 , wherein
(e) steps (a)-(c) include successively adding layers of metal powder joined by local directed energy such as direct laser metal sintering, selective laser metal melting, or electron beam melting; (f) step (d) is replaced by an injection molded ceramic core or stamped refractory metal negative for casting of features for at least one of the steps (a)-(c); and (g) step (d) further includes successively adding layers of metal powder to a partially cast component for construction of at least one of the steps (a)-(c).

17 . The method according to claim 16 , comprising additively manufacturing at least one core that provides a cavity having an airfoil shape corresponding to the airfoil, and the forming step includes casting the airfoil within the cavity.

18 . The method according to claim 15 , wherein the forming step includes casting a first airfoil portion, and additively manufacturing a second airfoil portion onto the first airfoil portion, the second airfoil portion providing the overhang.

19 . A method of manufacturing a gas turbine engine component, comprising the step of:
forming step a first airfoil portion, and additively manufacturing a second airfoil portion onto the first airfoil portion, the second airfoil portion including a recessed pocket that is circumscribed by an overhang, and at least one cooling groove provided by the overhang.

20 . The method according to claim 19 , wherein the first air foil portion is cast.

Assignees

• United Technologies Corp

Inventors

• Snyder Brooks E
• Slavens Thomas N

Classifications

• F01D5/20 Primary

  Specially-shaped blade tips to seal space between tips and stator {(F01D5/225 takes precedence)} · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• B22F10/28

  Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM] · CPC title

• B22F7/08

  with one or more parts not made from powder {(B22F7/062 takes precedence)} · CPC title

• F05D2230/22

  by sintering · CPC title