US 11,988,170 B2
Geared gas turbine engine
Alastair D Moore, Derby (GB); and Robert J Telling, Derby (GB)
Assigned to ROLLS-ROYCE plc, London (GB)
Filed by ROLLS-ROYCE PLC, London (GB)
Filed on Nov. 29, 2022, as Appl. No. 18/070,660.
Application 18/070,660 is a continuation of application No. 17/193,323, filed on Mar. 5, 2021, granted, now 11,542,890.
Application 17/193,323 is a continuation of application No. 16/398,707, filed on Apr. 30, 2019, granted, now 10,975,802, issued on Apr. 13, 2021.
Claims priority of application No. 1820945 (GB), filed on Dec. 21, 2018.
Prior Publication US 2023/0109777 A1, Apr. 13, 2023
This patent is subject to a terminal disclaimer.
Int. Cl. F02K 3/06 (2006.01); F02C 7/24 (2006.01)
CPC F02K 3/06 (2013.01) [F02C 7/24 (2013.01)] 17 Claims
OG exemplary drawing
 
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 extending from a hub, a diameter D of the fan being in a range of from 240 cm to 420 cm;
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, a gear ratio of the gearbox being in a range of from 2.5 to 5.0, wherein:
a ratio L/D is in a range of 0.33 to 0.48, L being an intake length defined as an axial distance between a leading edge of an intake of the engine and a leading edge of the plurality of fan blades at the hub, and D being the diameter of the fan;
the turbine drives the fan via the gearbox;
a minimum number of rotor blades in any single rotor stage of the turbine is in a range of from 60 to 140;
a take-off lateral reference point is defined as a point on a line parallel to and 450 m from a runway centre line where Effective Perceived Noise Level (EPNL) is a maximum during take-off of an aircraft to which the gas turbine engine is attached;
when the EPNL is a maximum at the take-off lateral reference point during the take-off, a relative Mach number of a tip of each of the plurality of fan blades does not exceed 1.09 M;
a bypass noise attenuation proportion L is defined as:

OG Complex Work Unit Math
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:

OG Complex Work Unit Math
where:
E is a total axial length of acoustic attenuation material provided to the intake; 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:

OG Complex Work Unit Math
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.