Gearboxes for aircraft gas turbine engines
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. A gearbox for an aircraft gas turbine engine includes: 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.
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 gearbox has a gear ratio in the range 3.2 to 4.5;
wherein the gas turbine engine is configured such that during operation a maximum specific load of each journal bearing is in the range of 5 MPa to 25 MPa.
2 . The gas turbine engine of claim 1 , wherein the planet carrier is connected to an output shaft to provide the output drive.
3 . The gas turbine engine of claim 1 , wherein the gearbox has a gear ratio in the range 3.6 to 4.5.
4 . The gas turbine engine of claim 1 , wherein the gearbox has a gear ratio in the range 3.6 to 4.2.
5 . The gas turbine engine of claim 1 , wherein the journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing.
6 . The gas turbine engine of claim 1 , configured such that, at cruise conditions the gas turbine engine has:
an overall pressure ratio of 40 to 60;
a bypass ratio between 12.5 and 16.5;
a specific thrust of less than 95 N kg-1s; and
a fan tip loading of 0.28 to 0.36.
7 . 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 55;
a bypass ratio between 13 and 16;
a specific thrust of 80 N kg-1s to 90 N kg-1s; and
a fan tip loading of 0.29 to 0.35.
8 . The gas turbine engine of claim 1 , wherein:
each one of the plurality of fan blades has a carbon-fibre body with a titanium leading edge.
9 . The gas turbine engine of claim 1 , wherein:
the diameter of the fan is in the range 220 cm to 300 cm; and
the rotational speed of the fan at cruise conditions is in the range 1700 rpm to 2500 rpm.
10 . The gas turbine engine of claim 1 , wherein:
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 0.31; and
the gas turbine engine comprises 18 fan blades.
11 . The gas turbine engine of claim 1 , wherein:
the fan diameter has an upper bound of 250 cm;
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 0.32;
the rotational speed of the fan at cruise conditions is less than 2500 rpm;
the fan has a fan tip loading at cruise conditions of 0.29 To 0.35; and
each one of the plurality of fan blades comprises a protective leading edge that is manufactured using a material that is better able to resist impact than the rest of the blade.
12 . The gas turbine engine of claim 1 , wherein:
the fan diameter is in the range 220 cm to 230 cm;
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 0.29;
the rotational speed of the fan at cruise conditions is less than 2500 rpm;
the fan has a fan tip loading at cruise conditions of 0.3 to 0.34;
each one of the plurality of fan blades has a carbon-fibre body with a titanium leading edge; and
the fan comprises exactly 16, 18, 20, 22, or 24 fan blades.
13 . The gas turbine engine of claim 1 , wherein:
a ratio of a length, L, of the internal and external sliding surfaces to the diameter, D, of the journal bearing is between 0.5 and 1.4; and
the journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing.
14 . The gas turbine engine of claim 1 , configured such that when it is operating at maximum take-off conditions, a Sommerfeld number of each journal bearing is greater than 4.
15 . The gas turbine engine of claim 1 , configured such that when it is operating at maximum take-off conditions, a Sommerfeld number of each journal bearing is between 10 and 16.
16 . The gas turbine engine of claim 1 , wherein an inefficiency of each journal bearing, defined as a percentage power loss with the aircraft gas turbine engine operating at maximum take-off conditions is less than 0.225% and no less than 0.1%.
17 . The gas turbine engine of claim 1 , wherein:
a ratio of a length, L, of the internal and external sliding surfaces to the diameter, D, of the journal bearing is between 0.5 and 1.4;
the journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing;
the gas turbine engine is configured such that when it is operating at maximum take-off conditions, a Sommerfeld number of each journal bearing is greater than 4; and
an inefficiency of each journal bearing, defined as a percentage power loss with the aircraft gas turbine engine operating at maximum take-off conditions is less than 0.225%.
18 . The gas turbine engine of claim 1 , wherein:
a ratio of a length, L, of the internal and external sliding surfaces to the diameter, D, of the journal bearing is between 0.5 and 1.4;
the journal bearing has a diametral clearance that is between 0.1% and 0.3% of a diameter of the journal bearing;
the gas turbine engine is configured such that when it is operating at maximum take-off conditions, a Sommerfeld number of each journal bearing is between 10 and 16;
an inefficiency of each journal bearing, defined as a percentage power loss with the aircraft gas turbine engine operating at maximum take-off conditions is less than 0.225% and no less than 0.1%;
the gearbox has a gear ratio greater than 3.6;
the planet carrier is connected to an output shaft to provide the output drive; and
the gas turbine engine is configured such that during operation a maximum specific load of each journal bearing is in the range of 9 MPa to 22 MPa.
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:
an inefficiency of each journal bearing, defined as a percentage power loss with the aircraft gas turbine engine operating at maximum take-off conditions is less than 0.225%; and
the gas turbine engine is configured such that during operation both of the following are 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 25 MPa.
20 . The gas turbine engine of claim 19 , wherein:
a diameter of the fan is greater than 360 cm; and
the gas turbine engine comprises no more than 14 fan blades.