Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes

Patent metadata

Field	Value
Publication number	US-2025382916-A1
Application number	US-202519248369-A
Country	US
Kind code	A1
Filing date	Jun 24, 2025
Priority date	Jun 14, 2024
Publication date	Dec 18, 2025
Grant date	—

Abstract

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.

First claim

We claim:

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 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 composite fan blades, a fan 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
-

rr

)

⁢

N
b

⁢

cos
⁡
(
γ
)

wherein c is a chord length of a composite fan blade of the plurality of composite fan blades, S is a span of the composite fan blade, rr is a radius ratio of the fan, γ is a stagger angle of the composite fan blade, and No is the number of the plurality of composite 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 composite fan blade trailing edge to an outlet guide vane leading edge measured parallel to a 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 core engine measured in pounds force at the max takeoff condition,
wherein

0.034

(

GR

1.5

)

(

D

1
.
5

⁢
6

T

)

1
.
5

⁢
3

>
5

,

wherein

0
.
0

⁢
1
⁢
5
⁢

(

GR

1.4

)

(

D

1
.
5

⁢
6

T

)

1
.
5

⁢
3

<
55

,
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 is 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 1.5 Nb to 3 Nb.

11 . The gas turbine engine of claim 1 , wherein Nv is 2.2 Nb to 2.6 Nb.

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 clai

Assignees

• Gen Electric

Inventors

• Miller Brandon Wayne
• Geertsema Egbert
• Sibbach Arthur W
• Hudecki Andrew
• Depuy Timothy Richard
• Schilling John C
• Worthoff Frank
• Nakano Tsuguji

Classifications

• F05D2260/961

  by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape · CPC title

• F05D2260/40311

  of the epicyclical, planetary or differential type · CPC title

• F02C7/36

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title

• F05D2300/603

  Composites; e.g. fibre-reinforced · CPC title

• F05D2260/96

  Preventing, counteracting or reducing vibration or noise · CPC title