# US 2014/0193238 A1 · Method for setting a gear ratio of a fan drive gear system of a gas turbine engine

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- Number: US 2014/0193238 A1
- Document ID: US20140193238A1
- Status: Published application
- Kind: A1
- Application number: 14/204,125
- Filed: 2014-03-11
- Published: 2014-07-10
- Earliest claimed: 2012-09-27
- Assignees: United Technologies Corporation
- Applicants: United Technologies Corporation
- Inventors: William G. Sheridan; Karl L. Hasel
- Web page: https://patentgrep.com/patent/US20140193238A1

## CPC

- F02C 7/32: Arrangement, mounting, or driving, of auxiliaries
- F02C 3/04: having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16)
- F02C 3/06: the compressor comprising only axial stages (F02C3/10 takes precedence)
- F02C 3/107: with two or more rotors connected by power transmission
- F02C 3/113: with variable power transmission between rotors
- F02C 7/00: Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants (controlling F02C9/00)
- F02C 7/36: 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)
- F16H 1/28: with gears having orbital motion
- F16H 1/36: with two central gears coupled by intermeshing orbital gears
- F05D 2210/12: gaseous, i.e. compressible
- F05D 2220/32: in gas turbines
- F05D 2220/36: specially adapted for the fan of turbofan engines
- F05D 2240/40: Use of a multiplicity of similar components
- F05D 2260/40311: of the epicyclical, planetary or differential type
- F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00)
- F16H: GEARING
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

A gas turbine engine includes a fan section including a fan that is rotatable about an axis. A speed reduction device is connected to the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A bypass ratio is greater than about 11.0.

## Claims

### Claim 1

A gas turbine engine comprising:
a fan section including a fan rotatable about an axis;
a speed reduction device connected to the fan, wherein the speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5, wherein a bypass ratio is greater than about 11.0.

### Claim 2 (depends on 1)

The gas turbine engine of claim 1, wherein the speed reduction device includes a star gear system gear ratio of at least 2.6.

### Claim 3 (depends on 2)

The gas turbine engine of claim 2, wherein the speed reduction device includes a system gear ratio less than or equal to 4.1.

### Claim 4 (depends on 3)

The gas turbine engine of claim 3, including the bypass ratio is less than about 22.0.

### Claim 5 (depends on 4)

The gas turbine engine of claim 4, wherein the fan blade tip speed of the fan section is greater than about 1000 ft/sec and less than about 1400 ft/sec.

### Claim 6 (depends on 1)

The gas turbine engine of claim 1, wherein the star system includes a sun gear, a plurality of star gears, a ring gear, and a carrier.

### Claim 7 (depends on 6)

The gas turbine engine of claim 6, wherein each of the plurality of star gears includes at least one bearing.

### Claim 8 (depends on 7)

The gas turbine engine of claim 7, wherein the carrier is fixed from rotation.

### Claim 9 (depends on 8)

The gas turbine engine of claim 8, wherein a low pressure turbine is mechanically attached to the sun gear.

### Claim 10 (depends on 9)

The gas turbine engine of claim 9, wherein a fan section is mechanically attached to the ring gear.

### Claim 11 (depends on 1)

The gas turbine engine of claim 1, wherein an input of the speed reduction device is rotatable in a first direction and an output of the speed reduction device is rotatable in a second direction opposite to the first direction.

### Claim 12 (depends on 11)

The gas turbine engine of claim 11, including a low pressure turbine section in communication with the speed reduction device, wherein the low pressure turbine section includes at least three stages.

### Claim 13

A method of improving performance of a gas turbine engine comprising:
determining fan tip speed boundary conditions for at least one fan blade of a fan section;
determining rotor boundary conditions for a rotor of a low pressure turbine;
utilizing stress level constraints in the rotor of the low pressure turbine and the at least one fan blade to determine if the rotary speed of the fan section and the low pressure turbine will meet a desired number of operating cycles, wherein a bypass ratio is greater than about 6.0, wherein a speed reduction device connects the fan section and the low pressure turbine and includes a star gear ratio of at least about 1.5.

### Claim 14 (depends on 13)

The method of claim 13, wherein the star gear ratio is less than about 4.1.

### Claim 15 (depends on 14)

The method of claim 14, wherein a fan pressure ratio is below 1.7.

### Claim 16 (depends on 15)

The method of claim 15, wherein a fan pressure ratio is below 1.48.

### Claim 17 (depends on 16)

The method of claim 16, wherein a bypass ratio is between about 11 and about 22.

### Claim 18 (depends on 17)

The method of claim 17, wherein a fan blade tip speed of the at least one fan blade is less than 1400 fps.

### Claim 19 (depends on 13)

The method of claim 13, wherein if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and the number of stages of the low pressure turbine is increased.

### Claim 20 (depends on 19)

The method of claim 19, wherein if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and an annular area of the low pressure turbine is increased.

### Claim 21

A fan drive gear module for a gas turbine engine comprising:
a star drive gear system with a speed reduction ratio of at least 1.5, wherein the star drive gear system is configured to drive a gas turbine engine with a bypass ratio greater than about 11.0.

### Claim 22 (depends on 21)

The fan drive gear module of claim 21, wherein the speed reduction ratio is greater than about 2.6 and less than or equal to 4.1.

### Claim 23 (depends on 21)

The fan drive gear module of claim 21, wherein the star drive gear system is configured to drive a fan section with a fan blade tip speed greater than about 1000 ft/sec and less than about 1400 ft/sec.

### Claim 24 (depends on 21)

The fan drive gear module of claim 21, wherein the bypass ratio is less than about 22.0.

### Claim 25

A method of designing a gas turbine engine comprising:
selecting fan tip speed boundary conditions for at least one fan blade of a fan section of a gas turbine engine;
selecting rotor boundary conditions for a rotor of a fan drive turbine of the gas turbine engine;
determining stress level constraints in the rotor of the fan drive turbine and the at least one fan blade to determine if the rotary speed of the fan section and the fan drive turbine will meet a desired number of operating cycles, wherein a bypass ratio is greater than about 6.0; and
connecting the fan section and the drive turbine through a speed reduction device that includes a speed reduction ratio of at least about 1.5.

### Claim 26 (depends on 25)

The method of claim 25, wherein a fan pressure ratio is below 1.7.

### Claim 27 (depends on 26)

The method of claim 26, wherein the speed reduction ratio is less than about 4.1.

### Claim 28 (depends on 27)

The method of claim 27, wherein the bypass ratio of the gas turbine engine is between about 11 and about 22.

### Claim 29 (depends on 25)

The method of claim 25, including lowering the speed reduction ratio of the speed reduction device and increasing a number of stages of the fan drive turbine responsive to determining that a stress level in the rotor or the at least one fan blade is outside a predefined number of operating cycles.

### Claim 30 (depends on 26)

The method of claim 26, including lowering the speed reduction ratio of the speed reduction device and increasing an annular area of the fan drive turbine responsive to determining that a stress level in the rotor or the at least one fan blade is outside a predefined number of operating cycles.