Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes
A gas turbine engine comprises a fan assembly, a core engine coupled to the fan assembly, a fan case housing the fan assembly and the core engine, a plurality of outlet guide vanes extending between the core engine and the fan case, and an acoustic spacing. The fan assembly comprises a plurality of fan blades that define a fan diameter and a blade effective acoustic length (BEAL). The acoustic spacing comprises a distance between the fan assembly and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. A gearbox assembly with an improved engine efficiency rating used in combination with the acoustic spacing provides an improved and balanced engine architecture.
1 . A gas turbine engine comprising:
a core engine comprising a low-pressure turbine;
a gearbox assembly including an input and an output, wherein the input is coupled to the low pressure turbine of the core engine and comprises a first rotational speed, wherein the output is coupled to a fan assembly and has a second rotational speed, and wherein a gear ratio (GR) of the first rotational speed to the second rotational speed is within a range of 2.5-5.0;
the fan assembly comprising a plurality of fan blades made from a composite material and having a fan blade diameter (D) measured in inches and a blade solidity that is greater than or equal to 0.8 and less than or equal 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 assembly;
a plurality of outlet guide vanes disposed aft of the fan assembly 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:
ASR
=
1
(
Nv
Nb
)
·
As
BEAL
wherein Nv is the number of the plurality of outlet guide vanes; and
a net thrust (T) of the gas turbine engine measured in pounds force at the max takeoff condition,
wherein
0
.
1
(
GR
1
.
5
)
(
D
1
.
5
6
T
)
1
.
5
3
N
2
>
3
,
wherein
GR
1
.
5
(
D
1.56
T
)
1.53
N
2
<
9
9
,
wherein N is a number of rotating blade stages of the low-pressure turbine, and
wherein the ASR is 1.5 to 16.0.
2 . The gas turbine engine of claim 1 , wherein the core engine comprises a high pressure compressor comprising 8-10 stages.
3 . The gas turbine engine of claim 1 , wherein the core engine comprises a high pressure compressor comprising 8-9 stages.
4 . The gas turbine engine of claim 1 , wherein:
the fan assembly comprises a fan pressure ratio of 1.30 to 1.35,
Nb is 14 to 20,
D is between 80 inches and 95 inches, and
the gas turbine engine comprises a bypass ratio between 12:1 and 17:1 at a takeoff condition.
5 . The gas turbine engine of claim 1 , wherein the fan assembly 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 20 to 22.
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 GR is 2 to 4.
15 . The gas turbine engine of claim 1 , wherein GR is 3.2 to 4.
16 . The gas turbine engine of claim 1 , wherein D is 80 inches to 95 inches.
17 . The gas turbine engine of claim 1 , wherein Nis three or more rotating blade stages.
18 . The gas turbine engine of claim 1 , wherein N is four or more rotating blade stages.
19 . The gas turbine engine of claim 1 , wherein Tis 12,000 pounds force to 35,000 pounds force.
20 . The gas turbine engine of claim 1 , wherein Tis 15,000 pounds force to 30,000 pounds force.