Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes
A gas turbine engine comprises a fan, a core engine coupled to the fan, a fan case housing the fan and the core engine, a plurality of outlet guide vanes extending between the core engine and the fan case, a blade effective acoustic length (BEAL), and an acoustic spacing. The fan comprises a fan disk, a fan diameter, and a plurality of fan blades. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. The fan case comprises an inlet, and in combination with the fan disk and fan diameter determines a disk spacing length and disk-to-blade diametric (DBD) ratio of the gas turbine engine.
1 . A gas turbine engine comprising:
a core engine comprising a low pressure turbine;
a gearbox assembly coupled to the low pressure turbine;
a fan coupled to the gearbox assembly, wherein the fan comprises:
a fan diameter;
a fan disk;
a plurality of fan blades made from a composite material; and
a blade solidity from 0.8 to 2.0;
a blade effective acoustic length (BEAL) defined as:
BEAL
=
2
c
2
S
(
1
-
r
r
)
N
b
cos
(
γ
)
wherein c is a chord length of a fan blade of the plurality of fan blades, 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;
a plurality of outlet guide vanes disposed aft of the fan and extending radially between the core engine and the fan case;
an acoustic spacing (As) from the fan blade trailing edge to an outlet guide vane leading edge measured parallel to a central longitudinal axis of the core engine;
an acoustic spacing ratio (ASR) defined as:
A
S
R
=
1
(
N
v
Nb
)
.
As
BEAL
wherein Nv is the number of the plurality of outlet guide vanes; and
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 the central longitudinal axis of the core engine,
wherein the ASR is 1.5 to 16.0, and the DBD ratio is 0.09 to 0.59.
2 . The gas turbine engine of claim 1 , wherein the DBD ratio is 0.15 to 0.35.
3 . The gas turbine engine of claim 1 , wherein the DBD ratio is 0.19 to 0.27.
4 . The gas turbine engine of claim 1 , wherein the fan comprises a fan pressure ratio of 1.25 to 1.45.
5 . The gas turbine engine of claim 1 , wherein the fan comprises a fan pressure ratio of 1.3 to 1.4.
6 . The gas turbine engine of claim 1 , wherein the gas turbine engine comprises a bypass ratio of 10:1 to 17:1 at a takeoff condition.
7 . The gas turbine engine of claim 1 , wherein the gas turbine engine comprises a bypass ratio of 12:1 to 15:1 at a takeoff condition.
8 . The gas turbine engine of claim 1 , wherein Nb is 14 to 26.
9 . The gas turbine engine of claim 1 , wherein:
Nb is 14 to 20,
the fan diameter is 95 inches to 105 inches,
the gas turbine engine comprises a bypass ratio of 12:1 to 17:1 at a takeoff condition, and
the fan comprises a fan pressure ratio between 1.30 and 1.35.
10 . The gas turbine engine of claim 1 , wherein Nv is 1.5Nb to 3Nb.
11 . The gas turbine engine of claim 1 , wherein Nv is 2.2Nb to 2.6Nb.
12 . The gas turbine engine of claim 1 , wherein the ASR is 4.0 to 14.0.
13 . The gas turbine engine of claim 1 , wherein the ASR is 6.6 to 13.5.
14 . The gas turbine engine of claim 1 , wherein the gearbox comprises a gear ratio of 2 to 4.
15 . The gas turbine engine of claim 1 , wherein the gearbox comprises a gear ratio of 3.2 to 4.
16 . The gas turbine engine of claim 1 , wherein the fan diameter is 80 inches to 95 inches.
17 . The gas turbine engine of claim 1 , wherein the low pressure turbine comprises at least three low pressure turbine stages.
18 . The gas turbine engine of claim 1 , wherein the low pressure turbine comprises at least four low pressure turbine stages.
19 . The gas turbine engine of claim 1 , wherein the radius ratio (rr) is 0.2 to 0.35.
20 . The gas turbine engine of claim 1 , wherein the stagger angle (γ) is 30° to 75°.