Geared gas turbine engine

1. A method of operating a gas turbine engine on 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 receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of 3.2 or higher,

wherein the method comprises operating the gas turbine engine to provide propulsion under cruise conditions such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range of from 0.75 to 1.0, and

wherein the jet velocity ratio, R J , is determined by the following formula:

R

J

=

V

B

⁢

C

B

V

C

⁢

C

C

⁢

η

L

⁢

P

⁢

T

⁢

η

F

where V B is the fully expanded first jet velocity, C B is a thrust coefficient of the bypass nozzle, V C is the fully expanded second jet velocity, C C is a thrust coefficient of the core exhaust nozzle, η LPT is an isentropic efficiency of a lowest pressure turbine of the engine core and η F is an isentropic efficiency of compression of air into the bypass duct by the fan.

2. The method of claim 1 , wherein the jet velocity ratio is within a range from 0.75 to 0.82 at cruise conditions.

3. The method of claim 1 wherein the gearbox is an epicyclic gearbox comprising an input sun gear connected to the core shaft, a plurality of planetary gears connected by a carrier arm and an outer annulus ring, the fan being connected to the carrier arm.

4. The method of claim 1 , wherein the fan has an outer diameter of between 220 cm and 390 cm.

5. The method of claim 4 , wherein the fan has an outer diameter of between 240 cm and 280 cm or between 330 cm and 380 cm.

6. The method of claim 1 , 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 rotate at a higher rotational speed than the first core shaft.

7. The method of claim 1 , wherein the gear ratio is between 3.2 and 3.8.

8. A method of operating a gas turbine engine on 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 receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of 3.4 or higher,

wherein the method comprises operating the gas turbine engine to provide propulsion under cruise conditions such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range of from 0.75 to 0.82, and

wherein the jet velocity ratio, R J , is determined by the following formula:

R

J

=

V

B

⁢

C

B

V

C

⁢

C

C

⁢

η

L

⁢

P

⁢

T

⁢

η

F

where V B is the fully expanded first jet velocity, C B is a thrust coefficient of the bypass nozzle, V C is the fully expanded second jet velocity, C C is a thrust coefficient of the core exhaust nozzle, η LPT is an isentropic efficiency of a lowest pressure turbine of the engine core and η F is an isentropic efficiency of compression of air into the bypass duct by the fan.

9. The method of claim 8 , wherein the gear ratio is within a range from 3.4 to 5.0.