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, Rj, 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, 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 ηF is an isentropic efficiency of a fan tip;
the jet velocity ratio, Rj, is between around 0.75 and 1.3 at cruise conditions;
a fan tip loading defined as dH/Utip 2 is between 0.28 and 0.35 at the cruise conditions, where dH is an enthalpy rise across the fan and Utip is a translational velocity of a leading edge of the fan tip; and
a temperature of a flow at an exit of the combustor, at a position immediately upstream of a first turbine vane, is at least 1600 K at the cruise conditions.
2 . The method of claim 1 , wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is between 1600 K and 1650 K at the cruise conditions.
3 . The method of claim 1 , wherein the temperature of the flow at the exit of the combustor, at a position immediately upstream of a first turbine vane, is between 1900 K and 2000 K at maximum take-off conditions.
4 . The method of claim 1 , wherein the fan tip loading at the cruise conditions is between 0.28 and 0.33.
5 . The method of claim 1 , wherein the fan tip loading at the cruise conditions is between 0.28 and 0.30.
6 . 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.
7 . The method of claim 1 , wherein the jet velocity ratio, R J , is between around 2 and 3 at idle conditions.
8 . The method of claim 1 , wherein a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is between 80 Nkg -1 s and 95 Nkg -1 s at the cruise conditions.
9 . The method of claim 1 , wherein a fan diameter is between 220 cm and 300 cm.
10 . The method of claim 9 , wherein a rotational speed of the fan at the cruise conditions is in the range of from 1700 rpm to 2500 rpm.
11 . The method of claim 1 , wherein a fan diameter is between 320 cm and 380 cm, and a rotational speed of the fan at the cruise conditions is in the range of from 1200 rpm to 2000 rpm.
12 . The method of claim 1 , wherein the gearbox has a planetary configuration and the jet velocity ratio, Rj, is below around 1.0 at the cruise conditions.
13 . The method of claim 1 , wherein the gearbox has a star configuration and the jet velocity ratio, Ri, is above around 1.0 at the cruise conditions.
14 . The method of claim 1 , wherein a gear ratio of the gearbox is above around 3.4, and the jet velocity ratio, Rj, is below around 1.0 at the cruise conditions.
15 . The method of claim 1 , wherein a gear ratio of the gearbox is below around 3.4, and the jet velocity ratio, Rj, is below between around 1.0 and 1.3 at the cruise conditions.
16 . The method of claim 1 , 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 the a flow through the core at the cruise conditions, is in a range of from 10.5 to 15.5.
17 . The method of claim 1 , wherein the bypass ratio is in a range from 12.5 to 13.5.
18 . The method of claim 1 , wherein 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 45 and 60 at the cruise conditions.
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 an aluminium based body with a titanium leading edge.