Gearboxes for aircraft gas turbine engines

1. A gas turbine engine for an aircraft, 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 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 journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing;

the gearbox has a gear ratio in the range 3 to 3.9;

the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip is in the range 0.25 to 0.33; and

the gas turbine engine is configured such that during operation a maximum specific load of each journal bearing is in the range 5 MPa to 20 MPa.

2. The gas turbine engine of claim 1 , wherein the diametral clearance is between 0.1% and 0.2% of the diameter of the journal bearing.

3. The gas turbine engine of claim 1 , wherein the diametral clearance is between 0.14% and 0.16% of the diameter of the journal bearing.

4. The gas turbine engine of claim 1 , wherein the diametral clearance is in the range 120 μm to 600 μm; and/or the diameter of the journal bearing is in the range 120 mm to 200 mm.

5. The gas turbine engine according to claim 1 , wherein:

the gear ratio is in the range 3.2 to 3.7; and/or

the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip is in the range 0.26 to 0.29.

6. The gas turbine engine according to claim 1 , wherein:

the diametral clearance is between 0.1% and 0.2% of the diameter of the journal bearing; and

the diametral clearance is in the range 120 μm to 600 μm; and

the gear ratio is in the range 3.1 to 3.5; and

the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip is in the range 0.27 to 0.31.

7. The gas turbine engine of claim 1 , wherein a ratio of a length of the internal and external sliding surfaces to the diameter of the journal bearing is between around 1.1 and 1.4.

8. The gas turbine engine of claim 1 , wherein:

a diameter of the fan is in the range 300 cm to 340 cm; and

a rotational speed of the fan at cruise conditions is in the range 1200 rpm to 2000 rpm.

9. The gas turbine engine of claim 1 , wherein, at cruise conditions, the gas turbine has any one or more of the following:

an overall pressure ratio in the range 45 to 50;

a bypass ratio of 10 to 11.5;

a specific thrust of 95 N kg −1 s to 100 N kg −1 s; and

a fan tip loading of 0.31 to 0.35.

10. The gas turbine engine of claim 1 , configured such that during operation the maximum specific load of each journal bearing is in the range 5 MPa to 10 MPa.

11. The gas turbine engine of claim 10 , wherein the gearbox has a star configuration and the gear ratio is in the range 3.1 to 3.4.

12. The gas turbine engine according to claim 1 , configured such that during operation a maximum operating sliding speed of each journal bearing is in the range 30 m/s to 50 m/s.

13. The gas turbine engine according to claim 1 , configured such that during operation a maximum operating sliding speed of each journal bearing is in the range 39 m/s to 50 m/s.

14. The gas turbine engine according to claim 1 , configured such that during operation:

the maximum specific load of each journal bearing is in the range 5 MPa to 10 MPa; and

a maximum operating sliding speed of each journal bearing is in the range 41 m/s to 50 m/s.

15. The gas turbine engine of claim 1 , configured such that, at cruise conditions the gas turbine engine has:

an overall pressure ratio of 45 to 50;

a bypass ratio between 11 and 12.5;

a specific thrust of 90 N kg-1s to 100 N kg-1s; and

a fan tip loading of 0.29 to 0.35, and

wherein the gas turbine engine is configured such that during operation:

the maximum specific load of each journal bearing is in the range 5 MPa to 10 MPa; and

a maximum operating sliding speed of each journal bearing is in the range 41 m/s to 50 m/s.

16. A gas turbine engine for an aircraft, 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 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 journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing;

the gearbox has a gear ratio in the range 3 to 3.9;

the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip is in the range 0.25 to 0.33; and

the gas turbine engine is configured such that during operation a maximum operating sliding speed of each journal bearing is in the range 30 m/s to 50 m/s.

17. The gas turbine engine of claim 16 , wherein:

the gas turbine engine is configured such that the maximum operating sliding speed of each journal bearing is in the range 41 m/s to 50 m/s; and/or

the diametral clearance is between 0.1% and 0.2% of the diameter of the journal bearing.

18. The gas turbine engine of claim 16 , wherein, at cruise conditions, the gas turbine has any one or more of the following:

an overall pressure ratio in the range 45 to 50;

a bypass ratio of 10 to 11.5;

a specific thrust of 95 N kg-1s to 100 N kg-1s; and

a fan tip loading of 0.31 to 0.35.

19. A gas turbine engine for an aircraft, 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 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 journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing;

the gearbox has a gear ratio of at least 4.2; and

the gas turbine engine is configured such that during operation at least one of the following is true: 1) a maximum operating sliding speed of each journal bearing is in the range 30 m/s to 50 m/s; and 2) a maximum specific load of each journal bearing is in the range 5 MPa to 20 MPa.

20. The gas turbine engine according to claim 19 , wherein:

a diameter of the fan is greater than 360 cm; and

a maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C. with the engine static is no greater than 170 kN.