Geared gas turbine engine with improved noise signature
A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine is provided that has high efficiency together with low noise, in particular from the fan and the turbine that drives the fan. Values are defined for a combined contribution of the turbine noise and the noise from the fan emanating from the front of the engine to the Effective Perceived Noise Level (EPNL) at take-off to be in the range of from 2 EPNdB and 15 EPNdB lower than the EPNL of the whole engine at take-off.
1 . A gas turbine engine for 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 bypass duct defined between an inner flow boundary formed by the engine core and an outer flow boundary formed by a nacelle; and
a gearbox that receives 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 gas turbine engine is configured such that, during a noise certification test of the aircraft, which is taking off from a runway and to which the gas turbine engine is attached, a combined contribution of (i) noise from the turbine and (ii) noise from the fan emanating from a front of the engine to a whole Effective Perceived Noise Level (EPNL) of the gas turbine engine at a take-off lateral reference point, defined as a point on a line parallel to and 450 m from a center line of the runway where the EPNL is a maximum during the take-off of the aircraft, is in a range of from 2 EPNdB and 15 EPNdB lower than the whole EPNL at the take-off lateral reference point;
a bypass noise attenuation proportion L is defined as:
L
=
J
+
H
2
G
where:
G is an axial length between tips of trailing edges of the fan blades and a trailing edge of the nacelle;
H is a total axial length of acoustic attenuation material provided to the outer flow boundary of the bypass duct over an axial extent between the tips of the trailing edges of the fan blades and the trailing edge of the nacelle; and
J is a total axial length of acoustic attenuation material provided to the inner flow boundary of the bypass duct over the axial extent between the tips of the trailing edges of the fan blades and the trailing edge of the nacelle;
an intake noise attenuation proportion K is defined as:
K
=
E
F
where:
E is a total axial length of acoustic attenuation material provided to an intake of the engine; and
F is an axial length of the intake;
a forward to rearward noise attenuation proportion M is in a range of from 0.8 to 2.5, where:
M
=
K
L
;
the intake noise attenuation proportion K is in a range of from 0.55 to 0.95;
a ratio of H/G is in a range of from 0.4 to 0.7; and
a ratio of J/G is in a range of from 0.4 to 0.7.
2 . The gas turbine engine according to claim 1 , wherein the combined contribution is in a range of from 5 EPNdB and 13 EPNdB lower than the whole EPNL of the gas turbine engine at the take-off lateral reference point.
3 . The gas turbine engine according to claim 1 , wherein
a core nozzle is provided downstream of a rearmost row of turbine blades in the turbine, the core nozzle being defined by a radially inner core nozzle boundary and a radially outer core nozzle boundary; and
at least one of the radially inner core nozzle boundary and the radially outer core nozzle boundary is provided with a noise attenuation liner.
4 . The gas turbine engine according to claim 1 , wherein a contribution of the turbine to the whole EPNL at the take-off lateral reference point is in a range of from 15 EPNdB and 40 EPNdB lower than a contribution of noise of the fan emanating from a rear of the engine to the whole EPNL at the take-off lateral reference point.
5 . The gas turbine engine according to claim 1 , wherein a contribution of the turbine to the whole EPNL at the take-off lateral reference point is in a range of from 25 EPNdB and 40 EPNdB lower than a contribution of noise of the fan emanating from a rear of the engine to the whole EPNL at the take-off lateral reference point.
6 . The gas turbine engine according to 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;
the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the first core shaft; and
noise of the first turbine contributes to the combined contribution.
7 . The gas turbine engine according to claim 1 , wherein a gear ratio of the gearbox is in a range of from 3 to 5.
8 . The gas turbine engine according to claim 1 , wherein a relative Mach number at a tip of each fan blade of the plurality of fan blades does not exceed 1.09 M at the take-off lateral reference point.
9 . The gas turbine engine according to claim 1 , wherein:
the turbine comprises at least two axially separated rotor stages; and
each and every one of the at least two axially separated rotor stages comprises in a range of from 60 to 140 rotor blades.
10 . The gas turbine engine according to claim 1 , wherein:
the turbine comprises at least two axially separated rotor stages; and
an average number of rotor blades in a rotor stage of the at least two axially separated rotor stages is in a range of from 65 to 120.
11 . The gas turbine engine according to claim 1 , wherein:
the turbine comprises at least two axially separated rotor stages; and
a number of rotor blades in a most axially rearward rotor stage of the at least two axially separated rotor stages is in a range of from 60 to 120.
12 . The gas turbine engine according to claim 1 , wherein a total number of turbine blades in the turbine is in a range of from 320 and 540.
13 . The gas turbine engine according to claim 1 , wherein a diameter of the fan is in a range of from 320 cm and 400 cm.
14 . The gas turbine engine according to claim 1 , wherein a bypass ratio at cruise conditions is in a range of from 12 to 18.
15 . A method of operating an aircraft comprising a gas turbine engine, 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 bypass duct defined between an inner flow boundary formed by the engine core and an outer flow boundary formed by a nacelle; and
a gearbox that receives 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, and the method comprising taking off from a runway during a noise certification test of the aircraft such that a combined contribution of (i) noise from the turbine and (ii) noise from the fan emanating from a front of the engine to a whole Effective Perceived Noise Level (EPNL) of the gas turbine engine at a take-off lateral reference point, defined as a point on a line parallel to and 450 m from a center line of the runway where the EPNL is a maximum during the take-off of the aircraft, is in a range of from 2 EPNdB and 15 EPNdB lower than the whole EPNL at the take-off lateral reference point, wherein:
a bypass noise attenuation proportion L is defined as:
L
=
J
+
H
2
G
where:
G is an axial length between tips of trailing edges of the fan blades and a trailing edge of the nacelle;
H is a total axial length of acoustic attenuation material provided to the outer flow boundary of the bypass duct over an axial extent between the tips of the trailing edges of the fan blades and the trailing edge of the nacelle; and
J is a total axial length of acoustic attenuation material provided to the inner flow boundary of the bypass duct over the axial extent between the tips of the trailing edges of the fan blades and the trailing edge of the nacelle;
an intake noise attenuation proportion K is defined as:
K
=
E
F
where:
E is a total axial length of acoustic attenuation material provided to an intake of the engine; and
F is an axial length of the intake;
a forward to rearward noise attenuation proportion M is in a range of from 0.8 to 2.5, where:
M
=
K
L
;
the intake noise attenuation proportion K is in a range of from 0.55 to 0.95;
a ratio of H/G is in a range of from 0.4 to 0.7; and
a ratio of J/G is in a range of from 0.4 to 0.7.
16 . The method according to claim 15 , wherein the combined contribution is in a range of from 5 EPNdB and 13 EPNdB lower than the whole EPNL of the gas turbine engine at the take-off lateral reference point.