Gearboxes for aircraft gas turbine engines

1. A gearbox for an aircraft gas turbine engine, 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, with the aircraft gas turbine engine operating at maximum take-off conditions, a minimum oil film thickness H min between the internal and external sliding surfaces is a function of a temperature T of oil flowing into the journal bearing, such that H min <B-AT, where A is 0.139 μm/° C. and B is 27.8 μm.

2. The gearbox of claim 1 , wherein the ring gear has a pitch circle diameter of around 550 mm or greater.

3. The gearbox of claim 1 , wherein H min is greater than 2.3 μm.

4. The gearbox of claim 1 , wherein H min >B′-A′T where A′ is 0.034 μm/° C. and B′ is 6.4 μm.

5. The gearbox of claim 1 , wherein H min >B″-A″T, where A″ is 0.117 μm/° C. and B″ is 22 μm.

6. The gearbox of claim 1 , wherein T is greater than around 60° C.

7. The gearbox ( 30 ) of claim 6 , wherein T is greater than around 100° C.

8. The gearbox ( 30 ) of claim 6 , wherein T is less than around 120° C.

9. The gearbox of claim 1 , wherein the gearbox has a gear ratio of 3.2 to 4.5 or 3.2 to 4.0.

10. The gearbox of claim 1 , wherein the gearbox is in a star configuration.

11. 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 according to claim 1 , the 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.

12. The gas turbine engine of claim 11 , wherein:

the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;

the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and

the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

13. The gas turbine engine according to claim 11 , wherein the gas turbine engine has:

a specific thrust from 70 to 90 N kg −1 ; and/or

a bypass ratio at cruise conditions of 12.5 to 18 or 13 to 16.

14. The gas turbine engine according to claim 11 , wherein:

the fan has a moment of inertia of between around 5.5×10 7 and 9×10 8 kg m 2 .

15. The gas turbine engine according to claim 11 , wherein the fan has a diameter on the order of 220 cm.

16. The gas turbine engine according to claim 11 , wherein the overall pressure ratio of the gas turbine engine at cruise conditions is 45 to 50.

17. The gas turbine engine according to claim 11 , wherein a fan tip loading of the fan at cruise conditions is 0.30 to 0.33.

18. The gas turbine engine according to claim 11 , wherein a hub-to-tip ratio of the fan is 0.25 to 0.29.

19. The gas turbine engine according to claim 11 , wherein the turbine entry temperature at cruise conditions is from 1600K to 1650K.

20. A method of operating a gas turbine engine according to claim 11 , the method comprising operating the aircraft gas turbine engine at maximum take-off conditions, a minimum oil film thickness H min between the internal and external sliding surfaces being a function of a temperature T of oil flowing into the journal bearing, such that H min <B-AT, where A is 0.139 μm/° C. and B is 27.8 μm.