Cooling channels for gas turbine engine component
A gas turbine engine component includes a wall portion and a leading edge cooling channel that extends through the wall portion. The leading edge cooling channel includes at least one first cooling passage separated from at least one serpentine cooling passage.
1. A gas turbine engine component comprising:
a wall portion; and
a leading edge cooling channel extending through the wall portion, the leading edge cooling channel including at least one first cooling passage separated from at least one serpentine cooling passage, wherein the at least one first cooling passage includes a pair of first cooling passages and the at least one serpentine cooling passage includes a pair of serpentine cooling passages, the pair of first cooling passages are L-shaped and at least partially surround a corresponding one of the pair of serpentine cooling passages.
2. The component of claim 1 , comprising a plurality of trip strips along at least one of a leading edge or a circumferential edge of the leading edge cooling channel.
3. The component of claim 1 , wherein the at least one first cooling passage includes a flow director in a bend of the at least one first cooling passage.
4. The component of claim 1 , wherein the at least one serpentine passage includes a serpentine channel rib extending at least partially along a length of the at least one serpentine passage along a mid-portion of the at least one serpentine passage.
5. The component of claim 1 , including a plurality of cooling channeling extending radially inward and upstream from a cooling cavity at least partially defined by the component through the component to the leading edge cooling channel.
6. The component of claim 1 , wherein the pair of first cooling passages are separated from each other by a central wall extending axially relative to a longitudinal axis of the gas turbine engine through the leading edge cooling passage.
7. The component of claim 6 , wherein the pair of first cooling passages and the pair of serpentine cooling passages mirror each other about the central wall and the pair of serpentine cooling passages include an axially extending portion along opposing edges of the component.
8. The component of claim 1 , wherein each of the pair of serpentine passages are located downstream of a corresponding one of the pair of first cooling passages and each of the pair of first cooling passages are separated from each of the pair of serpentine passages by an intermediate wall.
9. A gas turbine engine assembly comprising:
a blade outer air seal including a leading edge cooling channel extending through a wall portion, the leading edge cooling channel including at least one first cooling passage separated from at least one serpentine cooling passage; and
a seal at least partially axially aligned with the leading edge cooling channel and interfacing with the blade outer air seal, wherein the seal is a brush seal.
10. The assembly of claim 9 , including a plurality of cooling channeling extending radially inward and upstream from a cooling cavity at least partially defined by the blade outer air seal and through the blade outer air seal to the leading edge cooling channel.
11. The assembly of claim 9 , wherein the at least one first cooling passage includes a pair of first cooling passages and the at least one serpentine cooling passage includes a pair of serpentine cooling passages, the pair of first cooling passages are L-shaped and at least partially surround a corresponding one of the pair of serpentine cooling passages and a plurality of trip strips along at least one of a leading edge or a circumferential edge of the leading edge cooling channel.
12. The assembly of claim 9 , wherein the at least one first cooling passage includes a flow director located in a bend of the at least one first cooling passage.
13. The assembly of claim 9 , wherein the at least one serpentine passage includes a serpentine channel rib extending at least partially along a length of the at least one serpentine passage along a mid-portion of the at least one serpentine passage.
14. The assembly of claim 11 , wherein the pair of first cooling passages and the pair of serpentine cooling passages mirror each other about a central wall and the pair of serpentine cooling passages include an axially extending portion along opposing edges of the component.
15. A method of cooling a component in a gas turbine engine comprising:
directing a cooling fluid through a cooling channel in the component from a cooling cavity adjacent the component to a leading edge cooling channel; and
directing the cooling fluid outward towards opposing circumferential ends of the component through a first pair of cooling passages and a pair of serpentine passages, wherein the pair of first cooling passages are L-shaped and at least partially surround a corresponding one of the pair of serpentine cooling passages.
16. The method of claim 15 , wherein the first pair of cooling passages and the pair of serpentine passages mirror each other and the cooling fluid is directed radially inward and upstream through the cooling channel.
17. The method of claim 15 , comprising directing cooling fluid through each of the first pair of cooling passages with a flow director located in a bend in each of the first pair of cooling channels.
18. The method of claim 15 , comprising directing cooling fluid through each of the pair of serpentine passages with a serpentine channel rib extending at least partially along a length of each of the pair of serpentine passages along a mid-portion of each of the pair of serpentine passages and the pair of serpentine passages include a portion extending in an axial direction.