Turbine nozzle alignment with combustor of gas turbine engines
A gas turbine engine includes a compressor section, a combustion section defining a combustion chamber, and a turbine section disposed in serial flow order along a central axis of the gas turbine engine. The combustion section includes a plurality of fuel nozzles in fluid communication with the combustion chamber. The plurality of fuel nozzles define a fuel nozzle pitch extending between a fuel nozzle centerline of adjacent ones of the plurality of fuel nozzles. The turbine section includes a plurality of vanes and a plurality of rotor blades. The turbine section defines a clocking pitch fraction in degrees about the central axis and the clocking pitch fraction is defined between a peak temperature region and a midpoint of a vane pitch. The vane pitch is defined between the leading edge of adjacent ones of the plurality of vanes.
1 . A gas turbine engine, comprising:
a compressor section;
a combustion section defining a combustion chamber, the combustion section including a plurality of fuel nozzles in fluid communication with the combustion chamber, the plurality of fuel nozzles defining a fuel nozzle pitch extending between a respective fuel nozzle centerline of adjacent ones of the plurality of fuel nozzles; and
a turbine section, wherein the compressor section, the combustion section, and the turbine section are disposed in serial flow order along a longitudinal centerline axis of the gas turbine engine, the turbine section comprising:
a plurality of vanes, each vane of the plurality of vanes including a leading edge, a trailing edge, a pressure side extending between the leading edge and the trailing edge, and a suction side opposite the pressure side and extending between the leading edge and the trailing edge, and
a plurality of rotor blades configured to rotate at a rotor speed (Q) in radians per second about the longitudinal centerline axis, each rotor blade of the plurality of rotor blades including a blade leading edge, a blade trailing edge, a first blade side extending between the blade leading edge and the blade trailing edge, and a second blade side opposite the first blade side and extending between the blade leading edge and the blade trailing edge, each rotor blade of the plurality of rotor blades defining a blade chord extending between the blade leading edge and the blade trailing edge and a midspan radius (R b mid ) extending between the first blade side and the second blade side opposite the first blade side at a center of the blade chord, the midspan radius (R b mid ) extending perpendicular to the longitudinal centerline axis and measured at a midspan location of each rotor blade, wherein R b mid ·Ω is greater than or equal to 1,250 feet per second and less than or equal to 2,000 feet per second;
wherein each vane of the plurality of vanes defines a vane exit angle (α te ) extending from the trailing edge relative to the longitudinal centerline axis, the vane exit angle (α te ) is greater than or equal to 65 degrees and less than or equal to 80 degrees;
wherein the turbine section defines a clocking pitch fraction (θ C ) in degrees about the longitudinal centerline axis, the clocking pitch fraction (θ C ) defined between a peak temperature region and a midpoint of a vane pitch, the vane pitch defined between the respective leading edge of adjacent ones of the plurality of vanes, wherein the clocking pitch fraction (θ C ) comprises an angular offset of the peak temperature region from the midpoint of the vane pitch, and wherein the clocking pitch fraction (θ C ) is greater than or equal to 0.51 times the fuel nozzle pitch and less than or equal to 0.81 times the fuel nozzle pitch;
wherein the turbine section defines a suction side heat load (Λ S ) of the suction side of each of the plurality of vanes and a pressure side heat load (Λ P ) of the pressure side of each of the plurality of vanes;
wherein the suction side heat load (Λ S ) is equal to:
0.5
(
1
-
cos
(
2
π
(
θ
C
+
0
.
1
2
5
)
)
)
;
wherein the pressure side heat load (Λ P ) is equal to:
0.5
(
1
-
cos
(
2
π
(
θ
C
+
0
.
1
2
5
)
)
)
;
wherein the turbine section defines a turbine durability factor (τ) equal to:
1
Λ
s
+
20
[
(
1
-
[
0.7
Λ
p
+
0.3
Λ
s
]
4
)
·
sin
(
α
te
)
-
0.00025
·
R
b
mid
·
Ω
]
2
;
and wherein
1.
0
0
3
<
τ
τ
(
θ
c
=
0.5
)
<
1.27
1
.
2 . The gas turbine engine of claim 1 , wherein:
the turbine section comprises a plurality of vane segments; and
one vane of the plurality of vanes extends from each of the plurality of vane segments.
3 . The gas turbine engine of claim 1 , wherein a peak temperature angle between the peak temperature region and a respective one of the fuel nozzle centerlines is 0.3 times the fuel nozzle pitch.
4 . A gas turbine engine, comprising:
a compressor section;
a combustion section defining a combustion chamber, the combustion section including a plurality of fuel nozzles in fluid communication with the combustion chamber, the plurality of fuel nozzles defining a fuel nozzle pitch extending between a respective fuel nozzle centerline of adjacent ones of the plurality of fuel nozzles; and
a turbine section comprising a plurality of vane segments, wherein the compressor section, the combustion section, and the turbine section are disposed in serial flow order along a longitudinal centerline axis of the gas turbine engine, the turbine section comprising:
a plurality of vanes, each vane of the plurality of vanes including a leading edge, a trailing edge, a pressure side extending between the leading edge and the trailing edge, and a suction side opposite the pressure side and extending between the leading edge and the trailing edge, wherein two vanes of the plurality of vanes extend from each of the plurality of vane segments, and wherein the two vanes of each of the plurality of vane segments include a leading vane and a trailing vane, and
a plurality of rotor blades configured to rotate at a rotor speed (Ω) in radians per second about the longitudinal centerline axis, each rotor blade of the plurality of rotor blades including a blade leading edge, a blade trailing edge, a first blade side extending between the blade leading edge and the blade trailing edge, and a second blade side opposite the first blade side and extending between the blade leading edge and the blade trailing edge, each rotor blade of the plurality of rotor blades defining a blade chord extending between the blade leading edge and the blade trailing edge and a mid span radius (R b mid ) extending between the first blade side and the second blade side opposite the first blade side at a center of the blade chord, the midspan radius (R b mid ) extending perpendicular to the longitudinal centerline axis and measured at a midspan location of each rotor blade:
wherein R b_mid ·Ω is greater than or equal to 1.250 feet per second and less than or equal to 2,000 feet per second;
wherein each vane of the plurality of vanes defines a vane exit angle (α te ) extending from the trailing edge relative to the longitudinal centerline axis, the vane exit angle (α te ) is greater than or equal to 65 degrees and less than or equal to 80 degrees:
wherein the turbine section defines a clocking pitch fraction (θ C ) in degrees about the longitudinal centerline axis, the clocking pitch fraction (θ C ) defined between one of a plurality of peak temperature regions corresponding to a respective one of the plurality of fuel nozzles and a midpoint of a vane pitch, the vane pitch defined between the respective leading edge of adjacent ones of the plurality of vanes, wherein the clocking pitch fraction (θ C ) comprises an angular offset of the peak temperature region from the midpoint of the vane pitch, wherein the clocking pitch fraction (θ C ) is greater than or equal to 0.77 times the fuel nozzle pitch and less than or equal to 0.95 times the fuel nozzle pitch in degrees relative to the leading vanes, or wherein the clocking pitch fraction (θ C ) is greater than or equal to 0.28 times the fuel nozzle pitch and less than or equal to 0.45 times the fuel nozzle pitch in degrees relative to the trailing vanes;
wherein the turbine section defines:
a suction side heat load of each of the leading vanes (Λ LS ) is equal to:
0
.
5
(
1
-
cos
(
2
π
(
θ
c
+
0
.
8
3
)
)
)
,
a pressure side heat load of each of the leading vanes (Λ LP ) is equal to:
0
.
5
(
1
-
cos
(
2
π
(
θ
c
+
0
.
6
7
)
)
)
,
a suction side heat load of each of the trailing vanes (Λ TS ) is equal to:
0
.
5
(
1
-
cos
(
2
π
(
θ
c
+
0.33
)
)
)
,
and
a pressure side heat load of each of the trailing vanes (Λ TP ) is equal to:
0
.
5
(
1
-
cos
(
2
π
(
θ
c
+
0
.
1
7
)
)
)
,
wherein the turbine section defines a turbine durability factor (τ) including a leading vane turbine stage durability factor (τ LV ) of each of the leading vanes and a trailing vane turbine stage durability factor (τ TV ) of each of the trailing vanes;
wherein
τ
LV
=
1
Λ
LS
+
20
·
[
(
1
-
[
0.7
Λ
LP
+
0.3
Λ
LS
]
4
)
·
sin
(
α
te
)
-
0.00025
·
R
b
mid
·
Ω
]
2
;
wherein
τ
TV
=
1
Λ
TS
+
20
·
[
(
1
-
[
0.7
Λ
TP
+
0.3
Λ
TS
]
4
)
·
sin
(
α
te
)
-
0.00025
·
R
b
mid
·
Ω
]
2
;
wherein τ=max(τ LV ,τ TV ); and
wherein
1.
0
0
3
<
τ
τ
(
θ
c
=
0.5
)
<
1.271
.
5 . The gas turbine engine of claim 4 , wherein the pressure side of each of the trailing vanes is aligned with a respective one of the peak temperature regions.
6 . The gas turbine engine of claim 5 , wherein:
the trailing vane of the two vanes of each of the plurality of segments are clocked 0.28 to 0.45 times the fuel nozzle pitch in a direction opposite of rotation of the plurality of rotor blades from a respective one of the fuel nozzle centerlines extending through each of the plurality of fuel nozzles relative to the trailing vane.
7 . The gas turbine engine of claim 5 , wherein the trailing vane of the two vanes of each of the plurality of segments are clocked 0.02 to 0.15 times the fuel nozzle pitch in a direction opposite of rotation of the plurality of rotor blades from a respective one of the fuel nozzle centerlines extending through each of the plurality of fuel nozzles relative to the trailing vane.
8 . The gas turbine engine of claim 4 , wherein the pressure side of each of the leading vanes is aligned with a respective one of the peak temperature regions.
9 . The gas turbine engine of claim 8 , wherein:
the leading vane of the two vanes of each of the plurality of segments are clocked 0.77 to 0.95 times the fuel nozzle pitch in a direction opposite of rotation of the plurality of rotor blades from a respective one of the fuel nozzle centerlines extending through each of the plurality of fuel nozzles.