Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes
A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan blades feature a low aspect ratio, reducing blade count while maintaining thrust and efficiency. Efficiency is enhanced through a determined relationship between fan blade count, aspect ratio, and specific flow.
1 . A gas turbine engine comprising:
a core turbine engine comprising a low pressure turbine;
a gearbox assembly coupled to the low pressure turbine;
a fan coupled to the gearbox assembly and having a fan diameter, a fan area, and a plurality of composite fan blades with a blade solidity that is greater than or equal to 0.8 and less than or equal to 2.0, the plurality of composite fan blades having a chord length c1 at a 75% span position and a chord length c2 at a 50% span position;
a blade effective acoustic length (BEAL) defined as:
B
E
A
L
=
2
c
1
2
S
(
1
-
r
r
)
N
b
cos
(
γ
)
wherein, S is a span of the fan blade, rr is a radius ratio of the fan, γ is a stagger angle of the fan blade, and N b is the number of the plurality of fan blades;
a nacelle that includes a fan case that surrounds the fan, the fan case comprising an inlet disposed forward of the fan and an inlet length, wherein the inlet length is an axial distance between a leading edge of one of the plurality of fan blades and the inlet, as measured at a 75% span position of the fan blade;
a plurality of outlet guide vanes disposed aft of the fan and extending radially between the core turbine engine and the fan case;
an acoustic spacing from the fan blade trailing edge to an outlet guide vane leading edge;
an acoustic spacing ratio (ASR) defined as:
A
S
R
=
1
(
N
v
N
b
)
·
A
s
B
E
A
L
wherein As is the acoustic spacing and Nv is the number of the plurality of outlet guide vanes; and
wherein the ASR of the gas turbine engine is 1.5 to 16.0,
wherein the gas turbine engine is characterized by an improved fan assembly efficiency in which the following relationship applies:
1
5
<
N
b
*
A
R
*
50
S
F
<
6
0
wherein AR is a fan blade aspect ratio, and SF is a fan specific flow, and the fan blade aspect ratio (AR) is determined from S/c2, and the fan specific flow (SF) is determined from MF/FA, where MA is a mass flow of the fan and FA is a fan area,
wherein the number of composite fan blades (Nb) ranges from 12 to 22, and the fan blade aspect ratio (AR) ranges from 1.3 to 2.0.
2 . The gas turbine engine of claim 1 , wherein the number of composite fan blades (Nb) ranges from 12 to 18.
3 . The gas turbine engine of claim 1 , wherein the fan blade aspect ratio (AR) ranges from 1.5 to 1.9.
4 . The gas turbine engine of claim 1 , wherein the gearbox assembly has a gear ratio ranging from 2.7 to 4.0.
5 . The gas turbine engine of claim 1 , wherein the gearbox assembly has a gear ratio ranging from 3.25 to 3.75.
6 . The gas turbine engine of claim 1 , further comprising a fan pressure ratio from 1.25 to 1.45.
7 . The gas turbine engine of claim 1 , further comprising a disk-to-blade diametric (DBD) ratio defined as a ratio of a disk spacing length to the fan diameter, the disk spacing length being a distance between a forwardmost end of a fan disk and an intersection with the inlet taken along an engine centerline,
wherein the DBD ratio of the gas turbine engine is 0.15 to 0.35.
8 . The gas turbine engine of claim 7 , wherein the DBD ratio of the gas turbine engine is 0.19 to 0.27.
9 . The gas turbine engine of claim 1 , wherein a radius ratio (rr) of the fan is 0.25 to 0.30.
10 . The gas turbine engine of claim 1 , wherein the ratio Nv/Nb is 2.2 to 2.6.
11 . A gas turbine engine comprising:
a core turbine engine comprising a low pressure turbine;
a gearbox assembly coupled to the low pressure turbine, the gearbox assembly having a gear ratio within a range of 2.5 to 5.0;
a fan coupled to the gearbox assembly and having a fan diameter, a fan area, and a plurality of composite fan blades that have a chord length c1 at a 75% span position and a chord length c2 at a 50% span position;
a blade effective acoustic length (BEAL) defined as:
B
E
A
L
=
2
c
1
2
S
(
1
-
r
r
)
N
b
cos
(
γ
)
wherein S is a span of the fan blade, rr is a radius ratio of the fan, γ is a stagger angle of the fan blade, and N b is the number of the plurality of composite fan blades;
a nacelle that includes a fan case that surrounds the fan, the fan case comprising an inlet disposed forward of the fan and an inlet length, wherein the inlet length is an axial distance between a leading edge of one of the plurality of fan blades and the inlet, as measured at the 75% span position;
a plurality of outlet guide vanes disposed aft of the fan and extending radially between the core turbine engine and the fan case;
an acoustic spacing from the fan blade trailing edge to an outlet guide vane leading edge;
an acoustic spacing ratio (ASR) defined as:
A
S
R
=
1
(
N
v
N
b
)
·
A
s
B
E
A
L
wherein As is the acoustic spacing and Nv is the number of the plurality of outlet guide vanes;
wherein the ASR of the gas turbine engine is 4.0 to 16.0,
wherein AR is a fan blade aspect ratio, and SF is a fan specific flow, and the fan blade aspect ratio (AR) is determined from S/c2, and the fan specific flow (SF) is determined from MF/FA, where MA is a mass flow of the fan and FA is a fan area,
wherein the fan blade aspect ratio (AR) ranges from 1.3 to 2.0, and
wherein a ratio of the number of outlet guide vanes to the number of fan blades (Nv/Nb) is 2.0 to 2.5.
12 . The gas turbine engine of claim 11 , wherein the fan blade aspect ratio (AR) ranges from 1.5 to 2.0.
13 . The gas turbine engine of claim 11 , wherein the gearbox assembly has a gear ratio ranging from 3.2 to 4.0.
14 . The gas turbine engine of claim 11 , wherein the engine has a bypass ratio of 12:1 to 15:1.
15 . The gas turbine engine of claim 11 , wherein a fan blade diameter is 80 inches to 95 inches.
16 . The gas turbine engine of claim 11 , further comprising a disk-to-inlet (DIL) ratio defined as a ratio of the disk spacing length to the inlet length, wherein the DIL ratio is 0.40 to 0.70.
17 . The gas turbine engine of claim 16 , wherein the DIL ratio is 0.45 to 0.67.
18 . The gas turbine engine of claim 11 , wherein an inlet length ratio defined as a ratio of the inlet length to the fan diameter is 0.15 to 0.40.
19 . The gas turbine engine of claim 11 , wherein the engine is further characterized by an improved fan assembly efficiency parameter defined by the relationship:
19.5
<
(
N
b
×
A
R
×
50
)
/
S
F
<
4
7
.
2
.
20 . The gas turbine engine of claim 11 , wherein the gearbox assembly has a gear ratio ranging from 3.25 to 3.75.