Turbine engine with a blade assembly having a set of 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 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 Celsius (° C.) to 1120° C. and having a redline core speed (CS) in a range of 306 Hertz to 353 Hertz, 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 comprising:
a platform having an upper surface and a lower surface;
an airfoil extending from the upper surface of the platform, the airfoil having an exterior surface, the exterior surface defining a pressure side and a suction side, the airfoil including a leading-edge and a trailing-edge;
a shank extending from the lower surface, the shank having a base defining a base plane;
a leading-edge inlet passage located within the shank, the leading-edge inlet passage having:
a first cross-sectional area at a first plane, the first plane radially spaced 0.0167 meters from the base plane;
a second cross-sectional area at a second plane, the second plane radially spaced 0.0182 meters from the base plane, a first ratio of the second cross-sectional area and the first cross-sectional area defining a second normalized area (Â 2 ) in a range of 0.996 to 1.139;
a third cross-sectional area at a third plane, the third plane radially spaced 0.0212 meters from the base plane, a second ratio of the third cross-sectional area and the first cross-sectional area defining a third normalized area (Â 3 ) in a range of 1.485 to 1.620;
a fourth cross-sectional area at a fourth plane, the fourth plane radially spaced 0.0226 meters from the base plane, a third ratio of the fourth cross-sectional area and the first cross-sectional area defining a fourth normalized area (Â 4 ) in a range of 1.743 to 1.861;
a fifth cross-sectional area at a fifth plane, the fifth plane radially spaced 0.0277 meters from the base plane, a fourth ratio of the fifth cross-sectional area and the first cross-sectional area defining a fifth normalized area (Â 5 ) in a range of 1.155 to 1.262;
a sixth cross-sectional area at a sixth plane, the sixth plane radially spaced 0.0294 meters from the base plane, a fifth ratio of the sixth cross-sectional area and the first cross-sectional area defining a sixth normalized area (Â 6 ) in a range of 1.039 to 1.179; and
a seventh cross-sectional area at a seventh plane, the seventh plane radially spaced 0.0341 meters from the base plane, a sixth ratio of the seventh cross-sectional area and the first cross-sectional area defining a seventh normalized area (Â 7 ) in a range of 0.89 to 1.078; and
a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits including a group of suction side cooling conduits proximate to the suction side, the group of suction side cooling conduits having a first sum of areas at an eighth plane radially spaced 0.0400 meters from the base plane, the group of suction side cooling conduits having a second sum of areas at a ninth plane radially spaced 0.0476 meters from the base plane, an average of the first sum and the second sum defining a suction side cross-sectional area (SSA), wherein the suction side cross-sectional area (SSA) is from 0.000005 m 2 to 0.00000945 m 2 , and wherein,
20.927
≤
(
7
1
100
,
000
)
*
(
A
^
2
+
A
^
6
A
^
7
)
3
*
(
A
ˆ
4
A
^
5
)
A
^
3
*
(
0.005
m
2
S
S
A
(
m
2
)
)
*
(
EGT
(
°
C
.
)
221
(
°
C
.
)
)
*
(
CS
(
Hz
)
300
Hz
)
2
≤
1
7
1
.
4
2
7
.
2 . The gas turbine engine of claim 1 , further including a plurality of inlet passages located within the shank, the plurality of inlet passages including the leading-edge inlet passage.
3 . The gas turbine engine of claim 2 , wherein each of the inlet passages extends between the base and one or more of the plurality of cooling conduits.
4 . The gas turbine engine of claim 2 , wherein the plurality of inlet passages further includes a middle inlet passage and a trailing-edge inlet passage.
5 . The gas turbine engine of claim 1 , wherein the shank is configured as a dovetail.
6 . 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.
7 . The gas turbine engine of claim 1 , wherein the first cross-sectional area is equal to the seventh cross-sectional area.
8 . 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 Celsius (° C.) to 1120° C. and having a redline core speed (CS) in a range of 306 Hertz to 353 Hertz, 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;
an airfoil extending from the upper surface of the platform, the airfoil having an exterior surface, the exterior surface defining a pressure side and a suction side, the airfoil including a leading edge and a trailing edge;
a shank extending from the lower surface, the shank having a base defining a base plane;
a leading-edge inlet passage located within the shank, the leading-edge inlet passage having:
a first cross-sectional area at a first plane, the first plane radially spaced 0.0167 meters from the base plane;
a second cross-sectional area at a second plane, the second plane radially spaced 0.0182 meters from the base plane, a first ratio of the second cross-sectional area and the first cross-sectional area defining a second normalized area (Â 2 ) in a range of 0.996 to 1.139;
a third cross-sectional area at a third plane, the third plane radially spaced 0.0212 meters from the base plane, a second ratio of the third cross-sectional area and the first cross-sectional area defining a third normalized area (Â 3 ) in a range of 1.485 to 1.620;
a fourth cross-sectional area at a fourth plane, the fourth plane radially spaced 0.0226 meters from the base plane, a third ratio of the fourth cross-sectional area and the first cross-sectional area defining a fourth normalized area (Â 4 ) in a range of 1.743 to 1.861;
a fifth cross-sectional area at a fifth plane, the fifth plane radially spaced 0.0277 meters from the base plane, a fourth ratio of the fifth cross-sectional area and the first cross-sectional area defining a fifth normalized area (Â 5 ) in a range of 1.155 to 1.262;
a sixth cross-sectional area at a sixth plane, the sixth plane radially spaced 0.0294 meters from the base plane, a fifth ratio of the sixth cross-sectional area and the first cross-sectional area defining a sixth normalized area (Â 6 ) in a range of 1.039 to 1.179; and
a seventh cross-sectional area at a seventh plane, the seventh plane radially spaced 0.0341 meters from the base plane, a sixth ratio of the seventh cross-sectional area and the first cross-sectional area defining a seventh normalized area (Â 7 ) in a range of 0.89 to 1.078; and
a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits including a group of suction side cooling conduits proximate to the suction side, the group of suction side cooling conduits having a first sum of areas at an eighth plane radially spaced 0.0400 meters from the base plane, the group of suction side cooling conduits having a second sum of areas at a ninth plane radially spaced 0.0476 meters from the base plane, an average of the first sum and the second sum defining a suction side cross-sectional area (SSA), wherein the suction side cross-sectional area (SSA) is from 0.000005 m 2 to 0.00000945 m 2 , and wherein,
20.927
≤
(
7
1
100
,
000
)
*
(
A
^
2
+
A
^
6
A
^
7
)
3
*
(
A
ˆ
4
A
^
5
)
A
^
3
*
(
0.005
m
2
S
S
A
(
m
2
)
)
*
(
EGT
(
°
C
.
)
221
(
°
C
.
)
)
*
(
CS
(
Hz
)
300
Hz
)
2
≤
1
7
1
.
4
2
7
.
9 . The blade assembly of claim 8 , further including a plurality of inlet passages located within the shank, the plurality of inlet passages including the leading-edge inlet passage.
10 . The blade assembly of claim 9 , wherein each of the inlet passages extends between the base and one or more of the plurality of cooling conduits.
11 . The blade assembly of claim 9 , wherein the plurality of inlet passages further includes a middle inlet passage and a trailing-edge inlet passage.
12 . The blade assembly of claim 8 , wherein the shank is configured as a dovetail.
13 . The blade assembly of claim 8 , wherein the blade assembly is a stage one blade assembly of a high-pressure turbine of the gas turbine engine.
14 . The blade assembly of claim 8 , wherein the first cross-sectional area is equal to the seventh cross-sectional area.