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 combustor, 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 J , 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
L
P
T
η
F
where V B is a fully expanded first jet velocity, C B is a thrust coefficient of the bypass exhaust nozzle, V C is a fully expanded second jet velocity, C C 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 a fan tip;
the jet velocity ratio, R J , is between around 0.75 and 1.3 at maximum take-off conditions;
each fan blade has a radial span extending from a hub to a tip, and a ratio of a radius of each fan blade at its hub to a radius of each fan blade at its tip is in a rang of 0.25 to 0.4;
a bypass ratio, defined as a ratio of a mass flow rate of a flow through the bypass duct to a mass flow rate of a flow through the core at cruise conditions, is in a range of 10 to 17;
a fan tip loading defined as dH/U tip 2 is between 0.28 and 0.38 at cruise conditions, where dH is an enthalpy rise across the fan and U d is a translational velocity of the leading edge of a fan tip; and a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 80 Nkg-1s and 110 N kg −1 s at the cruise conditions.
2. The method of claim 1 , wherein: the bypass ratio is in a range of 10 to 12; the fan tip loading is between 0.29 and 0.35 at the cruise condition; and the specific thrust is between 90 Nkg −1 s and 105 Nkg −1 s at the cruise condition.
3. The method of claim 1 , wherein: the bypass ratio is in a range of 10.5 to 11.5; the fan tip loading is between 0.31 and 0.34 at cruise conditions; and the specific thrust is between 95 Nkg −1 s and 105 Nkg −1 s at the cruise conditions.
4. The method of claim 1 , wherein: the jet velocity ratio, t, is between around 0.8 and 1.0 at the maximum take-off conditions; the bypass ratio is in a range of 10 to 12; the fan tip loading is between 0.31 and 0.35 at the cruise conditions; and the specific thrust is between 90 Nkg −1 s and 100 Nkg −1 s at the cruise conditions.
5. The method of claim 1 , wherein: the jet velocity ratio, t, is between around 0.8 and 1.0 at the maximum take-off conditions; the ratio of the radius of each fan blade at its hub to the radius of each fan blade at its tip is in a range of 0.27 to 0.32; the bypass ratio is in a range of 10 to 12; the fan tip loading is between 0.3 and 0.35 at the cruise conditions; and the specific thrust is between 90 Nkg −1 s and 100 Nkg −1 s at the cruise conditions.
6. The method of claim 1 , wherein an area of a final rotor of the low pressure turbine is in a range of 0.25 m 2 to 0.38 m 2 .
7. The method of claim 6 , wherein a fan diameter is around 215 cm.
8. A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a combustor, 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 J , 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
L
P
T
η
F
where V B is a fully expanded first jet velocity, C B is a thrust coefficient of the bypass exhaust nozzle, V C is a fully expanded second jet velocity, C C 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 a fan tip;
the jet velocity ratio, R J , is between around 0.75 and 1.3 at maximum take-off conditions;
each fan blade has a radial span extending from a hub to a tip, and a ratio of a radius of each fan blade at its hub to a radius of each fan blade at its tip is in a rang of 0.25 to 0.4;
an overall pressure ratio defined as a ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor is between 40 and 65 at cruise conditions;
a par ratio of the gearbox is between 3 and 4; and
wherein a temperate of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is in a range of 1700K to 2000K at the maximum take-off conditions.
9. The method of claim 8 , wherein: the overall pressure ratio is between 40 and 50 at the cruise conditions; the gear ratio of the gearbox is between 3.1 and 3.4; and wherein the temperature of the flow at the exit of the combustor, at the position immediately upstream of the first turbine vane, is in a range of 1750K to 1950K at the maximum take-off conditions.
10. The method of claim 8 , wherein: the overall pressure ratio is between 45 and 50 at the cruise conditions; the gear ratio of the gearbox is between 3.2 and 3.4; and wherein the temperature of the flow at the exit of the combustor, at the position immediately upstream of the first turbine vane, is in a range of 1850K to 1950K at the maximum take-off conditions.
11. The method of claim 1 , wherein: the bypass ratio is in a range of 10 to 12; the fan tip loading is between 0.29 and 0.35 at the cruise conditions; the specific thrust is between 90 Nkg −1 s and 105 Nkg −1 s at the cruise conditions; an overall pressure ratio defined as a ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor is between 40 and 50 at the cruise condition; a gear ratio of the gearbox is between 3.0 and 3.5; and wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is in a range of 1750K to 1950K at the maximum take-off condition.
12. The method of claim 11 , wherein: the ratio of the radius of each fan blade at its hub to the radius of each fan blade at its tip is in a range of 0.27 to 0.32; and the jet velocity ratio, R J , is between around 0.8 and 1.0 at the maximum take-off conditions.
13. The method of claim 1 , wherein: the bypass ratio is in a rang of 10 to 11.5; the fan tip loading is between 0.31 and 0.34 at the cruise conditions; the specific thrust is between 90 Nkg −1 s and 100 Nkg −1 s at the cruise conditions; an overall pressure ratio defined as a ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor is between 40 and 50 at the cruise conditions; a par ratio of the gearbox is between 3.2 and 3.4; and wherein a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is in a rang of 1850K to 1950K at the maximum take-off conditions.
14. The method of claim 13 , wherein: the ratio of the radius of each fan blade at its hub to the radius of each fan blade at its tip is in a range of 0.28 to 0.31; and the jet velocity ratio, j, is between around 0.8 and 1.0 at the maximum take-off conditions.
15. A method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising:
an engine core comprising a turbine, a combustor, 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 J , 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
L
P
T
η
F
where V B is a fully expanded first jet velocity, C B is a thrust coefficient of the bypass exhaust nozzle, V C is a fully expanded second jet velocity, C C 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 a fan tip;
the jet velocity ratio, t, is between around 0.75 and 1.3 at maximum take-off conditions; and each fan blade has a radial span extending from a hub to a tip, and a ratio of a radius of each fan blade at its hub to a radius of each fan blade at its tip is in a range of 0.25 to 0.4.
16. The method of claim 15 , wherein:
the jet velocity ratio, R J , is between around 2 and 3 at idle conditions; and
the jet velocity ratio, R J , is between around 0.75 and 1.3 at cruise conditions.
17. The method of claim 16 , wherein a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 90 Nkg −1 s and 105 Nkg −1 s at the cruise conditions.
18. The method of claim 17 , wherein a fan tip loading defined as dH/U is between 0.29 and 0.35 at the cruise conditions, where dH is an enthalpy rise across the fan and U d is a translational velocity of a leading edge of a fan tip.
19. The method of claim 18 , wherein the ratio of the radius of each fan blade at its hub to the radius of each fan blade at its tip is in a range of 0.27 to 0.32.
20. The method of claim 19 , wherein a bypass ratio, defined as a ratio of a mass flow rate of a flow through the bypass duct to a mass flow rate of a flow through the core at the cruise conditions, is in a range of 10 to 12.