Gas turbine engine turbine blade tip cooling

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 having a radially inwardly extending lip that provides an interior perimeter of the pocket, and at least one cooling groove provided by the overhang, wherein the structure is an airfoil.

2. The component according to claim 1 , wherein cooling fluid exits through a continuous channel into the recessed pocket.

3. The component according to claim 1 , wherein cooling fluid 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 airfoil includes a cast first portion, and a second portion is secured to the first portion, the second portion providing the overhang.

6. The component according to claim 5 , wherein the second portion is additively manufactured.

7. The component according to claim 1 , wherein the groove is provided between the overhang and the end wall, the groove bounded by the lip.

8. The component according to claim 7 , wherein the overhang substantially encloses the groove and provides an exit that fluidly interconnects the groove with the pocket.

9. The component according to claim 8 , wherein the exit is provided radially between the lip and the end wall.

10. The component according to claim 1 , wherein the pocket is teardrop-shaped.

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

12. 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 a negative for casting of features for at least one of the steps (a)-(c) and using an additive manufacturing process to create the negative for casting of features for at least one of the steps (a)-(c); and

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.

13. The method according to claim 12 , 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) comprises using 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).

14. The method according to claim 13 , 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.

15. 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.

16. The method according to claim 15 , wherein the first airfoil portion is cast.