Turbine engine with a blade assembly having cooling holes
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. Each blade including cooling holes arranged in an overlapping region of the blade.
1 . A turbine engine, comprising:
an engine core configured to generate an exhaust gas temperature (EGT) in a range of 980.1 and 1059.4 degrees Celsius (° C.) and a rotational speed (N2) in a range of 12511 to 13980.5 RPM, 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 configured to rotate about the engine centerline at the rotational speed (N2), wherein a radial axis is perpendicular to the engine centerline and aligned with a height dimension of the blade assembly, the blade assembly comprising:
a platform defining a platform plane;
an airfoil extending from the platform, the airfoil including:
a pressure side and a suction side and an exterior surface;
a leading edge and a trailing edge;
a root;
a tip defining a tip plane;
a cooling conduit;
a plurality of cooling holes in an overlapping region of the pressure side of the blade;
wherein an airfoil height (H airfoil ) is defined as a radial distance measured from the platform plane to the tip plane, wherein the airfoil height is in a range of 73.7 to 86.4 millimeters;
wherein the overlapping region of the pressure side of the blade is defined by:
a mid-region, wherein the mid-region is a region of the pressure side of the airfoil between a first radial plane and a second radial plane, wherein the first radial plane is arranged parallel to the tip plane and located at a first radial distance from the engine centerline, wherein the second radial plane is arranged parallel to the tip plane and located at a second radial distance from the engine centerline, and wherein the first radial distance is 0.369 meters and the second radial distance is 0.353 meters; and
an aft region, wherein the a blade curve is defined as a line of intersection between the exterior surface of the pressure side and the second radial plane, wherein the blade curve extends between the leading edge and the trailing edge, wherein a curve point is defined at a distance of 68% of a length of the blade curve from the leading edge along the blade curve; wherein an aft plane includes the curve point, is perpendicular to the second radial plane, and is perpendicular to the pressure side of the blade curve at the curve point, wherein the aft region is a region of the pressure side of the blade that is located between the aft plane and the trailing edge;
wherein each cooling hole in the plurality of cooling holes includes a cooling passage, each cooling passage extending between an inlet fluidly coupled to the cooling conduit and an outlet open to the exterior surface, the cooling passage defining a passage centerline;
wherein each cooling hole in the plurality of cooling holes includes a radial hole angle (RHA) defined between the second radial plane and the respective passage centerline;
wherein a radial hole angle average (RHA avg ) is defined as an average of the radial hole angles (RHAs) for all cooling holes in the overlapping region, wherein the radial hole angle average (RHA avg ) is in the range of 21.4 to 63.7 degrees; and
wherein a minimum radial hole angle (RHA min ) is defined as a smallest radial hole angle of all the radial hole angles in the plurality of cooling holes, wherein the minimum radial hole angle (RHA min ) is in the range of 20.1 to 44.9 degrees;
wherein the exhaust gas temperature (EGT), the rotational speed (N2), the airfoil height (H airfoil ), and the radial hole angle average (RHA avg ), define a first value (EQ1) by the following expression:
EQ
1
=
(
(
RHA
avg
15
°
)
+
3
.
7
5
(
H
airfoil
(
mm
)
73.7
mm
)
3
)
(
1
.
7
×
EGT
(
°
C
.
)
1000
°
C
.
)
(
1.2
×
13
,
000
RPM
N
2
(
RPM
)
)
,
wherein the exhaust gas temperature (EGT), the rotational speed (N2), and the minimum radial hole angle (RHA min ), define a second value (EQ2) by the following expression:
EQ
2
=
(
RHA
min
30
°
)
(
EGT
(
°
C
.
)
1000
°
C
.
)
(
13
,
000
RPM
N
2
(
RPM
)
)
,
and
wherein the first value (EQ1) is in the range of 9.87 to 22.7, and the second value (EQ2) is in the range of 0.61 to 1.64.
2 . The turbine engine of claim 1 , wherein the tip plane and the platform plane are parallel to one another.
3 . The turbine engine of claim 2 , wherein the tip plane, the platform plane, the first radial plane, and the second radial plane are perpendicular to the radial axis.
4 . The turbine engine of claim 1 , wherein 0% of the blade curve is defined at the leading edge and 100% of the blade curve is defined at the trailing edge.
5 . The turbine engine of claim 1 , wherein the passage centerline is a longitudinal axis of a respective one of the plurality of cooling holes.
6 . The turbine engine of claim 5 , wherein the passage centerline extends through a center of an inlet and a center of an outlet of a respective one of the plurality of cooling holes.
7 . The turbine engine of claim 1 , wherein the radial hole angle (RHA) is an angle between the passage centerline and a projection of the passage centerline onto the second radial plane.
8 . The turbine engine of claim 1 , wherein the platform plane includes a platform point on an exterior surface of the platform and is perpendicular to the radial axis.
9 . The turbine engine of claim 8 , wherein the platform point is on the exterior surface of the platform that is configured to confront a mainstream airflow of the turbine engine.
10 . The turbine engine of claim 9 , wherein the platform plane is adjacent the root of the airfoil.
11 . The turbine engine of claim 1 , wherein the tip plane is defined as a plane including a tip point on the tip of the airfoil and is perpendicular to the radial axis.
12 . The turbine engine of claim 1 , wherein the radial axis intersects the center of mass of the blade assembly or the airfoil.