Geared gas turbine engine
A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.
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;
a nacelle surrounding the engine core and defining a bypass duct and a bypass exhaust nozzle; and
a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft,
wherein the method comprises operating the gas turbine engine to provide propulsion such that a jet velocity ratio, R 3 , of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core is defined as:
R
J
=
V
B
C
B
V
C
C
C
η
LPT
η
F
where V B is a fully expanded first jet velocity, C B is a thrust coefficient of the bypass exhaust nozzle, Vc is a fully expanded second jet velocity, Cc is a thrust coefficient of the engine 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 the fan tip, the jet velocity ratio, R j , is between around 0.75 and 1.3 at maximum take-off conditions, a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is less than 90 Nkg −1 s at cruise conditions, and the fan diameter is between 220 cm and 390 cm.
2. The method of claim 1 , wherein the jet velocity ratio, R J , is between around 0.8 and 1.0 at maximum take-off conditions.
3. The method of claim 1 , wherein the jet velocity ratio, R J , is between around 2 and 3 at idle conditions.
4. The method of claim 1 , wherein the jet velocity ratio, R J , is between around 0.75 and around 1.3 at cruise conditions.
5. The method of claim 1 , wherein the specific thrust is between 80 Nkg −1 s and 90 Nkg −1 s at cruise conditions.
6. The method of claim 1 , wherein the fan diameter is between 220 cm and 300 cm.
7. The method of claim 6 , wherein a rotational speed of the fan at cruise conditions is in the range of from 1700 rpm to 2500 rpm.
8. The method of claim 1 , wherein the fan diameter is between 220 cm and 280 cm.
9. The method of claim 1 , wherein the fan diameter is between 320 cm and 380 cm.
10. The method of claim 9 , wherein a rotational speed of the fan at cruise conditions is in the range of from 1200 rpm to 2000 rpm.
11. The method of claim 9 , wherein a rotational speed of the fan at cruise conditions is in the range of from 1400 rpm to 1800 rpm.
12. The method of claim 1 , wherein the fan diameter is between 330 cm and 360 cm.
13. The method of claim 1 , wherein the gearbox has a planetary configuration and the jet velocity ratio, R J , is below around 1.0 at cruise conditions.
14. The method of claim 1 , wherein the gearbox has a star configuration and the jet velocity ratio, R J , is above around 1.0 at cruise conditions.
15. The method of claim 1 , wherein a gear ratio of the gearbox is above around 3.4, and the jet velocity ratio, R J , is below around 1.0 at cruise conditions.
16. The method of claim 1 , wherein a gear ratio of the gearbox is below around 3.4, and the jet velocity ratio, R J , is below between around 1.0 and 1.3 at cruise conditions.
17. The method of claim 1 , wherein a bypass ratio, defined as the ratio of a mass flow rate of a flow through the bypass duct to a mass flow rate of a flow through the engine core at cruise conditions, is in a range of from 12.5 to 16.5.
18. The method of claim 17 , wherein the bypass ratio is in a range from 13 to 15.
19. The method of claim 1 , wherein the fan comprises 16, 18 or 20 fan blades.
20. The method of claim 1 , wherein each fan blade comprises a carbon-fibre or aluminium based body with a protective leading edge.