Using inserts to balance heat transfer and stress in high temperature alloys

1. A method for forming a gas turbine engine component comprising the steps of:

(a) forming a block of a high temperature alloy material;

(b) machining an external surface of the block to form an external surface of a gas turbine engine component;

(c) machining at least a first cooling passage and a second cooling passage separated from the first cooling passage by a wall within the component that are open to at least one end of the component; and

(d) forming at least first and second inserts with a heat transfer feature and installing the first insert within the first cooling passage and installing the second insert within the second cooling passage, and including forming the first and second inserts to have an internal surface and an external surface that is spaced apart from internal wall surfaces that form the first and second cooling passages, and including forming the heat transfer feature to comprise a plurality of protruding portions on the insert that do not contact the internal wall surfaces.

2. The method according to claim 1 wherein the high temperature alloy material can withstand operating temperatures within a range of 2400-2700 degrees Fahrenheit.

3. The method according to claim 1 wherein the high temperature alloy material comprises molybdenum or a monolithic ceramic material.

4. The method according to claim 1 wherein step (a) includes forging the block as a single piece structure.

5. The method according to claim 4 including performing step (b) prior to step (c).

6. The method according to claim 4 including performing step (c) prior to step (b).

7. The method according to claim 1 wherein step (c) includes machining the cooling passages to have smooth walls that comprise the internal wall surfaces.

8. The method according to claim 1 including (e) welding each insert to the airfoil body.

9. The method according to claim 1 wherein step (d) includes forming each insert as a hollow body that is open to at least one end of the insert, and with each insert having the heat transfer feature formed on at least one of the external surface and the internal surface of the insert.

10. The method according to claim 1 including forming the at least one heat transfer feature as a plurality of pins extending outwardly from a surface of the insert into the cooling passage.

11. The method according to claim 1 including forming the protruding portions as a plurality of rounded protrusions extending outwardly from a surface of the insert into the cooling passage.

12. The method according to claim 1 including forming protruding portions as a plurality of trip strips extending outwardly from a surface of the insert into the cooling passage.

13. The method according to claim 1 including forming the protruding portions as a plurality of dimples forming recesses on a surface of the insert.

14. A component for a gas turbine engine comprising:

a body formed from a high temperature alloy material, the body extending between an outer surface and an inner surface spaced radially inward of the outer surface;

at least a first cooling passage and a second cooling passage separated from the first cooling passage by a wall, wherein the at least first and second cooling passages are formed in the body, and are open to at least one of the outer and inner surfaces, and wherein the first and second cooling passages have smooth walls extending in a radial direction, and are spaced apart from each other in an axial direction that is transverse to the radial direction;

at least a first insert and a second insert, each with a heat transfer feature, and wherein the first insert is positioned within the first cooling passage and the second insert is positioned within the second cooling passage, and wherein the first and second inserts have an internal surface and an external surface that is spaced apart the smooth walls, and wherein the heat transfer feature comprises a plurality of protruding portions on the insert that do not contact the smooth walls; and

a cover attached to the body over an open end of the cooling passage to enclose the insert within the body.

15. The component according to claim 14 wherein the high temperature alloy material can withstand operating temperatures within a range of 2400-2700 degrees Fahrenheit.

16. The component according to claim 14 wherein the high temperature alloy material comprises molybdenum, a monolithic ceramic material, or a ceramic matrix composite material.

17. The component according to claim 14 wherein the protruding portions comprise at least one of

a plurality of pins extending outwardly from the insert into the cooling passage,

a plurality of rounded protrusions extending outwardly from the insert into the cooling passage,

a plurality of trip strips extending outwardly from the insert into the cooling passage, or

a plurality of dimples forming recesses on the insert.

18. The component according to claim 14 wherein the body comprises one of a gas turbine engine vane, blade, BOAS, or combustor panel.

19. The component according to claim 14 wherein the body comprises a forged material that can withstand operating temperatures within a range of 2400-2700 degrees Fahrenheit.

20. The component according to claim 14 wherein each insert comprises a hollow body that extends continuously from the inner surface to the outer surface.

21. The method according to claim 1 including enclosing the insert within the cooling passage by placing a cover on the at least one end of the component to cover an open end of the cooling passage.

22. The method according to claim 1 including

forming the insert by bending sheet metal to form a hollow body, or

forming the insert by an additive manufacturing process.