Granted Patent B2

US 12,560,125 · App. 19/177,403 · Granted Feb 24, 2026

Inventors: Brandon Wayne Miller (Evendale, OH); Egbert Geertsema (Evendale, OH); Arthur W. Sibbach (Boxford, MA); Andrew Hudecki (Evendale, OH); Timothy Richard DePuy (Evendale, OH); John C. Schilling (Evendale, OH); Frank Worthoff (Evendale, OH); Tsuguji Nakano (Evendale, OH)

Assignee: General Electric Company

F02C7/24F02C7/36F02K3/06F05D2260/40311F05D2260/961

Abstract

Claims (85)

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 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, wherein the gas turbine engine further comprises an inlet-to-nacelle (ITN) ratio defined as a ratio of the inlet length to a maximum diameter of the nacelle, and wherein the ITN ratio is 0.23 to 0.35.

Assignments (1)

Continuity (2)

References Cited (110)