Cooling passages for gas turbine engine component
A gas turbine engine component includes a wall portion that includes a first side and a second opposite side. A plurality of passages extends between the first side of the wall portion and the second side of the wall portion and includes a plurality of inlets located on the first side of the wall portion. A plurality of outlets are located on a second side of the wall portion. The plurality of outlets include a first plurality of outlets located on a first side of the plurality of inlets and a second plurality of outlets located on a second side of the plurality of inlets.
1. A gas turbine engine component comprising:
a wall portion including a first side and a second opposite side; and
a plurality of passages extending between the first side of the wall portion and the second opposite side of the wall portion comprising:
a plurality of inlets located on the first side of the wall portion, wherein each of the plurality of inlets are axially aligned; and
a plurality of outlets located on the second opposite side of the wall portion, wherein the plurality of outlets include a first plurality of outlets located upstream of the plurality of inlets and a second plurality of outlets located downstream of the plurality of inlets.
2. The component of claim 1 , wherein the first plurality of outlets and the second plurality of outlets are positioned in an alternating pattern.
3. The component of claim 1 , wherein the plurality of passages are directed in a circumferential direction.
4. The component of claim 1 , wherein the plurality of passages are directed in a radial direction.
5. The component of claim 1 , wherein the first plurality of outlets are located in a first axial position and the second plurality of outlets are located in a second axial position.
6. The component of claim 5 , wherein the second axial position extends in a second direction parallel to a first direction of the first axial position.
7. The component of claim 1 , wherein the gas turbine engine component is at least one of a blade outer air seal, a vane, a blade, or a combustor liner.
8. A gas turbine engine comprising:
at least one blade outer air seal including a wall portion having a first side and a second side; and
a plurality of passages extending between the first side of the wall portion and the second side of the wall portion comprising:
a plurality of inlets located on the first side of the wall portion at a first axial position, wherein each of the plurality of inlets are axially aligned; and
a plurality of outlets located on the second side of the wall portion, wherein the plurality of outlets include a first plurality of outlets located upstream from a second plurality of outlets.
9. The gas turbine engine of claim 8 , wherein the first plurality of outlets and the second plurality of outlets are positioned in an alternating pattern.
10. The gas turbine engine of claim 8 , wherein the first plurality of outlets are located upstream of the plurality of inlets and the second plurality of outlets are located downstream of the plurality of inlets.
11. The gas turbine engine of claim 8 , wherein the plurality of passages are directed in a circumferential direction.
12. The gas turbine engine of claim 8 , wherein the plurality of inlets and the plurality of outlets are elliptical.
13. The gas turbine engine of claim 8 , wherein a diameter of at least one of the plurality of passages increases from the first side of the wall portion to the second side of the wall portion.
14. The gas turbine engine of claim 8 , wherein the first plurality of outlets are located in a first axial position and the second plurality of outlets are located in a second axial position and the second axial position extends in a second direction parallel to a first direction of the first axial position.
15. A method of cooling a gas turbine engine component comprising:
directing a cooling fluid into a first plurality of linearly arranged passages, each passage of the first plurality of linearly arranged passages including an inlet on a first side of the gas turbine engine component, wherein each of the inlets of the first plurality of linearly arranged passages are axially aligned; and
film cooling a second side of the gas turbine engine component with the cooling fluid exiting the first plurality of linearly arranged passages through a first plurality of outlets having a first linear arrangement and a second plurality of outlets having a second linear arrangement spaced from the first linear arrangement, wherein the first plurality of outlets are upstream of the second plurality of outlets.
16. The method of claim 15 , wherein the first linear arrangement is parallel to the second linear arrangement.
17. The method of claim 15 , wherein the first plurality of outlets are located upstream of said each of the inlets of the first plurality of linearly arranged passages and the second plurality of outlets are located downstream of said each of the inlets of the first plurality of linearly arranged passages.
18. The method of claim 15 , wherein the gas turbine engine component is at least one of a blade outer air seal, a vane, a blade, or a combustor liner.