Gearboxes for aircraft gas turbine engines

Patent metadata

Field	Value
Publication number	US-12385441-B2
Application number	US-202318239314-A
Country	US
Kind code	B2
Filing date	Aug 29, 2023
Priority date	Apr 6, 2020
Publication date	Aug 12, 2025
Grant date	Aug 12, 2025

Abstract

Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines, including a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

First claim

The invention claimed is:

1. A method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
a gearbox that is configured to receive an input from the core shaft and provide an output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox comprising:
a sun gear;
a plurality of planet gears surrounding and engaged with the sun gear; and
a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin,
wherein the internal or external sliding surface of the journal bearing has a surface coating,
wherein a fan tip loading is defined as dH/Utip2, where dH is a enthalpy rise across the fan and Utip is a velocity of the fan tip, and the fan tip loading at cruise conditions is in a range of from 0.30 Jkg−1/(ms−1)2 to 0.35 Jkg−1/(ms−1)2,
and wherein the method comprises operating the engine at maximum take-off conditions such that a specific loading multiplied by a sliding speed for each journal bearing is greater than around 240 MPa m/s.

2. The method of claim 1 , wherein the method comprises operating the engine at maximum take-off conditions such that a specific loading multiplied by a sliding speed for each journal bearing is less than around 720 MPa m/s.

3. The method of claim 1 , wherein a thickness of the surface coating is between around 40 and around 200 micrometres.

4. The method of claim 1 , wherein the fan tip loading at cruise conditions is in a range of from 0.31 Jkg−1/(ms−1)2 to 0.35 Jkg−1/(ms−1)2.

5. The method of claim 1 , wherein the fan tip loading at cruise conditions is in a range of from 0.31 Jkg−1/(ms−1)2 to 0.34 Jkg−1/(ms−1)2.

6. The method of claim 1 , wherein the gear ratio of the gearbox is in a range from 3.2 to 3.7.

7. The method of claim 1 , wherein a bypass ratio of the gas turbine engine at cruise conditions is in a range from 10.5 to 12.0.

8. A method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
a gearbox that is configured to receive an input from the core shaft and provide an output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox comprising:
a sun gear;
a plurality of planet gears surrounding and engaged with the sun gear; and
a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin,
wherein the internal or external sliding surface of the journal bearing has a surface coating,
wherein the method comprises operating the engine at maximum take-off conditions such that:
a specific loading multiplied by a sliding speed for each journal bearing is less than around 720 MPa m/s; and
and a temperature of a flow at exit of the combustor is in a range from 1850 K to 1950 K.

9. The method of claim 8 , wherein the method comprises operating the engine at maximum take-off conditions such that a specific loading multiplied by a sliding speed for each journal bearing is greater than around 240 MPa m/s.

10. The method of claim 8 , wherein a diameter of the fan is in a range from 220 cm to 240 cm.

11. The method of claim 8 , wherein the fan comprises either 22 or 24 fan blades.

12. The method of claim 8 , wherein a diameter of the fan is in a range from 220 cm to 240 cm and the rotational speed of the fan at cruise conditions is in the range of from 1700 rpm to 2500 rpm.

13. A method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
a gearbox that is configured to receive an input from the core shaft and provide an output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox comprising:
a sun gear;
a plurality of planet gears surrounding and engaged with the sun gear; and
a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin,
wherein the internal or external sliding surface of the journal bearing has a surface coating,
wherein an overall pressure ratio of a gas turbine engine, defined as a ratio of a stagnation pressure upstream of the fan to a stagnation pressure at the exit of a highest pressure compressor, is in a range from 45 to 55 at cruise conditions,
and wherein the method comprises operating the engine at maximum take-off conditions such that a specific loading multiplied by a sliding speed for each journal bearing is greater than around 240 MPa m/s.

14. The method of claim 13 , wherein the method comprises operating the engine at maximum take-off conditions such that a specific loading multiplied by a sliding speed for each journal bearing is less than around 720 MPa m/s.

15. The method of claim 13 , wherein a maximum operating sliding speed of each journal bearing is in a range from 35 m/s to 49 m/s.

16. The method of claim 13 , wherein the gas turbine engine has a specific thrust from 90 to 100 N kg−1 at cruise conditions.

17. The method of claim 16 , wherein the gas turbine engine has a specific thrust from 95 to 100 N kg−1 at cruise conditions.

18. The method of claim 16 , wherein the overall pressure ratio of the gas turbine engine at cruise conditions is in a range 45 to 50.

19. The method of claim 16 , wherein the bypass ratio of the gas turbine engine at cruise conditions is in the range 10.5 to 11.5.

20. The method of claim 19 , wherein a fan tip loading is defined as dH/Utip2, where dH is the enthalpy rise across the fan and Utip is the velocity of the fan tip, and the fan tip loading at cruise conditions is in a range of from 0.30 Jkg−1/(ms−1)2 to 0.33 Jkg−1/(ms−1)2.

Assignees

• Rolls Royce Plc

Inventors

• Spruce Mark

Classifications

• F16H2057/085

  Bearings for orbital gears · CPC title

• F16H2057/02043

  for vehicle transmissions · CPC title

• F16H57/08 Primary

  of gearings with members having orbital motion · CPC title

• F16H57/02

  Gearboxes; Mounting gearing therein · CPC title

• F16H1/28

  with gears having orbital motion · CPC title