Turbine engine with a blade assembly having cooling conduits
A gas turbine engine having a blade assembly with a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet passages to provide cooling fluid to the cooling conduits in the airfoil. The cooling fluid is vented through a plurality of cooling holes along the trailing edge of the airfoil. The blade assembly has specific geometries that improve durability.
1. A gas turbine engine, comprising:
an engine core configured to generate a redline exhaust gas temperature (EGT) in a range of 988 degrees Celsius (C) to 1120° C., the engine core extending along an engine centerline and including:
a compressor section;
a combustor; and
a turbine section, the turbine section including a blade assembly rotatable about the engine centerline, the blade assembly including:
a platform having an upper surface and a lower surface, the platform having a stator rotor seal with an upper edge having a radius of curvature defined as a stator rotor seal radius (SRSR), wherein the stator rotor seal radius (SRSR) is 0.224 to 0.239 meters;
an airfoil coupled to the upper surface of the platform, the airfoil having an outer wall defining an exterior surface, the exterior surface defining a pressure side and a suction side, the outer wall extending between a leading-edge and a trailing-edge;
a shank coupled to the lower surface, the shank having a base defining a base plane; and
a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits including:
a first cooling conduit located closest to the trailing-edge and defining a trailing-edge area (TEA), wherein the trailing-edge area (TEA) is from 0.0000056 square-meters (m 2 ) to 0.00001 m 2 ; and
a second cooling conduit next to the first cooling conduit and defining a secondary area (SA), wherein the secondary area (SA) is from 0.000002 m 2 to 0.0000048 m 2 , and wherein,
43.951
≤
[
100
*
(
TEA
1
(
m
2
)
0.005
m
2
)
-
2
(
SA
(
m
2
)
0.005
m
2
)
2
]
[
0.17
*
(
Redline
EGT
(
°
C
.
)
500
°
C
.
)
(
SRSR
(
m
)
1
m
)
]
≤
9
7
6
.
3
8
8
.
2. The gas turbine engine of claim 1 , wherein the trailing-edge area (TEA) is an average of a first cross-sectional area of the first cooling conduit located at a first plane extending through the airfoil at a first radial distance of 0.04002 meters measured from the base plane of the shank and a second cross-sectional area of the first cooling conduit located at a second plane extending through the airfoil at a second radial distance of 0.04764 meters measured from the base plane of the shank.
3. The gas turbine engine of claim 2 , wherein the secondary area (SA) is an average of a third cross-sectional area of the second cooling conduit located at the first plane and a fourth cross-sectional area of the second cooling conduit located at the second plane.
4. The gas turbine engine of claim 1 , wherein the first cooling conduit and the second cooling conduit are sized to provide the outer wall with a thickness that provides sufficient durability to the blade assembly.
5. The gas turbine engine of claim 1 , wherein the shank includes a plurality of inlet passages fluidly coupled to the plurality of cooling conduits.
6. The gas turbine engine of claim 5 , wherein each of the inlet passages extends between the base and one or more of the cooling conduits.
7. The gas turbine engine of claim 6 , wherein the plurality of inlet passages includes a leading-edge inlet passage, a middle inlet passage, and a trailing-edge inlet passage.
8. The gas turbine engine of claim 7 , wherein the first cooling conduit is fluidly coupled to the trailing-edge inlet passage, and the second cooling conduit is fluidly coupled to the middle inlet passage.
9. The gas turbine engine of claim 1 , wherein the blade assembly is a stage one blade assembly of a high-pressure turbine of the turbine section.
10. The gas turbine engine of claim 1 , wherein the shank is configured as a dovetail.
11. A blade assembly for a gas turbine engine having an engine core configured to generate a redline exhaust gas temperature (EGT) in a range of 988 degrees Celsius (° C.) to 1120° C., the blade assembly to be connected to the engine core and rotatable about an engine centerline of the engine core, the blade assembly comprising:
a platform having an upper surface and a lower surface, the platform having a stator rotor seal with an upper edge having a radius of curvature defined as a stator rotor seal radius (SRSR), wherein the stator rotor seal radius (SRSR) is 0.224 to 0.239 meters;
an airfoil coupled to the upper surface of the platform, the airfoil having an outer wall defining an exterior surface, the exterior surface defining a pressure side and a suction side, the outer wall extending between a leading-edge and a trailing-edge;
a shank coupled to the lower surface, the shank having a base defining a base plane; and
a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits including:
a first cooling conduit located closest to the trailing-edge and defining a trailing-edge area (TEA), wherein the trailing-edge area (TEA) is from 0.0000056 square-meters (m 2 ) to 0.00001 m 2 ; and
a second cooling conduit next to the first cooling conduit and defining a secondary area (SA), wherein the secondary area (SA) is from 0.000002 m 2 to 0.0000048 m 2 , and wherein,
43.951
≤
[
100
*
(
TEA
1
(
m
2
)
0.005
m
2
)
-
2
(
SA
(
m
2
)
0.005
m
2
)
2
]
[
0.17
*
(
Redline
EGT
(
°
C
.
)
500
°
C
.
)
(
SRSR
(
m
)
1
m
)
]
≤
9
7
6
.
3
8
8
.
12. The blade assembly of claim 11 , wherein the trailing-edge area (TEA) is an average of a first cross-sectional area of the first cooling conduit located at a first plane extending through the airfoil at a first radial distance of 0.04002 meters measured from the base plane of the shank and a second cross-sectional area of the first cooling conduit located at a second plane extending through the airfoil at a second radial distance of 0.04764 meters measured from the base plane of the shank.
13. The blade assembly of claim 12 , wherein the secondary area (SA) is an average of a third cross-sectional area of the second cooling conduit located at the first plane and a fourth cross-sectional area of the second cooling conduit located at the second plane.
14. The blade assembly of claim 11 , wherein the first cooling conduit and the second cooling conduit are sized to provide the outer wall with a thickness that provides sufficient durability to the blade assembly.
15. The blade assembly of claim 11 , wherein the shank includes a plurality of inlet passages fluidly coupled to the plurality of cooling conduits.
16. The blade assembly of claim 15 , wherein each of the inlet passages extends between the base and one or more of the cooling conduits.
17. The blade assembly of claim 16 , wherein the plurality of inlet passages includes a leading-edge inlet passage, a middle inlet passage, and a trailing-edge inlet passage.
18. The blade assembly of claim 17 , wherein the first cooling conduit is fluidly coupled to the trailing-edge inlet passage, and the second cooling conduit is fluidly coupled to the middle inlet passage.
19. The blade assembly of claim 11 , wherein the blade assembly is a stage one blade assembly of a high pressure turbine.
20. The blade assembly of claim 11 , wherein the shank is configured as a dovetail.