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

Patent metadata

Field	Value
Publication number	US-12560125-B2
Application number	US-202519177403-A
Country	US
Kind code	B2
Filing date	Apr 11, 2025
Priority date	Jun 14, 2024
Publication date	Feb 24, 2026
Grant date	Feb 24, 2026

Abstract

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 comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. 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. A gearbox assembly with 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 4.1-7.0; the fan assembly comprising a plurality of fan blades made from a composite material and having 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 assembly, γ is a stagger angle of the fan blade, and N b is the number 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 outlet guide vanes;
a gearbox efficiency rating of 0.10-0.70, wherein the gearbox efficiency rating equals

Q
⁡
(

D

1
.
5

⁢
6

T

)

1
.
5

⁢
3

,

 wherein Q is a gearbox oil flow rate at an inlet of the gearbox assembly measured in gallons per minute at a max takeoff condition, wherein D is a diameter of the fan blades measured in inches, and wherein T is a net thrust of the gas turbine engine measured in pounds force at the max takeoff condition,
wherein the ASR of the gas turbine engine is 1.5 to 16.0.

2 . The gas turbine engine of claim 1 , wherein the ASR range of 1.5 to 16.0 mitigates an increase in acoustic noise that results from increased propulsive power by the gearbox efficiency rating of 0.10-0.70.

3 . The gas turbine engine of claim 1 , wherein the gearbox efficiency rating is 0.10-0.62.

4 . The gas turbine engine of claim 1 , wherein the gearbox efficiency rating is 0.21-0.51.

5 . The gas turbine engine of claim 1 , wherein the gear ratio is within a range of 4.1-5.1.

6 . The gas turbine engine of claim 1 , wherein Q is within a range of 6-36 gallons per minute.

7 . The gas turbine engine of claim 1 , wherein T is within a range of 12,000-30,000 pounds force.

8 . The gas turbine engine of claim 1 , wherein the gearbox assembly is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is the input, and wherein the ring gear is the output.

9 . The gas turbine engine of claim 1 , wherein the gearbox assembly is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is the input, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is the output.

10 . The gas turbine engine of claim 1 , further comprising a fan pressure ratio from 1.30 to 1.55.

11 . The gas turbine engine of claim 1 , wherein the ASR is 4.0 to 14.0.

12 . The gas turbine engine of claim 1 , wherein the ASR is 6.6 to 13.5.

13 . The gas turbine engine of claim 12 , further comprising a disk-to-nacelle diametric (DND) 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 DND ratio of the gas turbine engine is 0.07 to 0.47.

14 . The gas turbine engine of claim 13 , wherein the DND ratio of the gas turbine engine is 0.15 to 0.35.

15 . The gas turbine engine of claim 13 , wherein the DND ratio of the gas turbine engine is 0.15 to 0.25.

16 . The gas turbine engine of claim 1 , wherein the fan case comprises an inlet disposed forward of the fan assembly 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, and wherein the gas turbine engine further comprises a disk-to-inlet length (DIL) ratio defined as a ratio of a disk spacing length to the inlet length, 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 DIL ratio of the gas turbine engine is 0.30 to 0.80.

17 . The gas turbine engine of claim 16 , wherein the DIL ratio of the gas turbine engine is 0.30 to 0.70.

18 . The gas turbine engine of claim 16 , wherein the DIL ratio of the gas turbine engine is 0.49 to 0.65.

19 . The gas turbine engine of claim 16 , further comprising a fan pressure ratio from 1.25 to 1.45.

20 . The gas turbine engine of claim 1 , wherein the fan case comprises an inlet disposed forward of the fan assembly and an inlet leng

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

• 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

• 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

• F02K3/06

  with front fan · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title