Large-scale bypass fan configuration for turbine engine core and bypass flows

Patent metadata

Field	Value
Publication number	US-2020200080-A1
Application number	US-201916398742-A
Country	US
Kind code	A1
Filing date	Apr 30, 2019
Priority date	Dec 21, 2018
Publication date	Jun 25, 2020
Grant date	—

Abstract

A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:

bypass



exhaust



nozzle



pressure



ratio

core



exhaust



nozzle



pressure



ratio

is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

First claim

1 . A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a core exhaust nozzle having a core exhaust nozzle exit, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core exhaust nozzle exit; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, the bypass duct comprising a bypass exhaust nozzle having a bypass exhaust nozzle exit, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass exhaust nozzle exit; wherein a bypass to core ratio of:

bypass



exhaust



nozzle



pressure



ratio

core



exhaust



nozzle



pressure



ratio

is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

2 . The gas turbine engine of claim 1 , wherein the bypass to core ratio is in the range from 1.10 to 2.00, under aircraft cruise conditions.

3 . The gas turbine engine of claim 1 , wherein the bypass to core ratio is above 1.15 under aircraft cruise conditions.

4 . The gas turbine engine of claim 1 , wherein the fan has a fan tip radius, and either:
(i) the fan tip radius is in the range from 110 cm to 150 cm, and the bypass to core ratio is in the range from 1.1 to 1.4; or (ii) the fan tip radius is in the range from 155 cm to 200 cm, and the bypass to core ratio is in the range from 1.3 to 1.6.

5 . The gas turbine engine of claim 1 , wherein the total pressure at the bypass exhaust nozzle exit is determined at an exit plane of the bypass exhaust nozzle, the exit plane extending from a rearmost point of the nacelle towards a centreline of the gas turbine engine.

6 . The gas turbine engine of claim 1 , wherein the engine core comprises a casing, and wherein the total pressure at the core exhaust nozzle exit is determined at an exit plane of the core exhaust nozzle, the exit plane extending from a rearmost point of the casing towards a centreline of the gas turbine engine.

7 . The gas turbine engine of claim 1 , wherein a flow area of the bypass exhaust nozzle is in the range from 2 m 2 to 6 m 2 .

8 . The gas turbine engine of claim 7 , wherein the fan has a fan tip radius, and
the fan tip radius is in the range from 155 cm to 200 cm, and the flow area of the bypass exhaust nozzle is in the range from 4.5 m 2 to 5.8 m 2 .

9 . The gas turbine engine of claim 1 , wherein a flow area of the core exhaust nozzle is in the range from 0.4 m 2 to 1.3 m 2 .

10 . The gas turbine engine of a claim 9 , wherein the fan has a fan tip radius, and either:
(i) the fan tip radius is in the range from 110 cm to 150 cm, and the flow area of the core exhaust nozzle is in the range from 0.4 m 2 to 0.6 m 2 ; or (ii) the fan tip radius is in the range from 155 cm to 200 cm, and the flow area of the core exhaust nozzle is in the range from 0.6 m 2 to 1.3 m 2 .

11 . The gas turbine engine of claim 1 , wherein at least one of the bypass exhaust nozzle and the core exhaust nozzle is a convergent nozzle.

12 . The gas turbine engine of claim 1 , wherein a bypass ratio defined as a ratio of mass flow rate of flow through the bypass duct to mass flow rate of flow through the engine core at cruise conditions is in the range of from 11 to 20.

13 . The gas turbine engine of claim 12 , wherein the fan has a fan tip radius, and either:
(i) the fan tip radius is in the range from 110 cm to 150 cm, and the bypass ratio is in the range from 13 to 16; or (ii) the fan tip radius is in the range from 155 cm to 200 cm, and the bypass ratio is in the range from 13 to 18.

14 . The gas turbine engine of claim 1 , further comprising a gearbox connected between the core shaft and the fan, the gearbox receiving an input from the core shaft and providing an output to drive the fan at a lower rotational speed than the core shaft, and wherein, optionally, the gearbox has a gear ratio in the range of from 3.2 to 3.8.

15 . The gas turbine engine of claim 1 , wherein a fan tip radius of the fan, measured between a centreline of the gas turbine engine and an outermost tip of each of the plurality of fan blades at its leading edge, is within a range of 110 cm to 200 cm.

16 . The gas turbine engine of 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; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

17 . (canceled)

Assignees

• Rolls Royce Plc

Inventors

• Stretton Richard G
• Willmot Michael C
• Grech Nicholas

Classifications

• F02K3/06 Primary

  with front fan · CPC title

• F02K1/00

  Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto (rocket nozzles F02K9/97) · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F04D29/388

  characterised by construction · CPC title

• F04D29/541

  Specially adapted for elastic fluid pumps (F04D29/56 takes precedence) · CPC title