Combustor tile for a combustor of a gas turbine engine

Patent metadata

Field	Value
Publication number	US-2020348024-A1
Application number	US-202016825067-A
Country	US
Kind code	A1
Filing date	Mar 20, 2020
Priority date	Mar 21, 2019
Publication date	Nov 5, 2020
Grant date	—

Abstract

An annular combustor for a gas turbine engine including inner and outer combustor walls, wherein each wall defines an annulus and the inner wall is radially inward of the outer wall. The combustor includes a primary zone where the inner and outer combustor walls converge in a downstream direction, and a secondary zone downstream of the primary zone. In the secondary zone, the inner and outer walls are arranged to converge at a different rate to the primary zone, are non-convergent or are divergent in a downstream direction, a rate of change of radial width of the combustor is different in the zones. A transition is provided from the primary zone to the secondary zone. A plurality of combustor cooling tiles lines the inner and outer walls. One or more of the tiles are arranged to extend from the primary to secondary zone and across the transition from the zones.

First claim

We claim:

1 . An annular combustor for a gas turbine engine, the combustor comprising:
an inner combustor wall and an outer combustor wall, the inner and outer combustor walls each define an annulus and the inner combustor wall is radially inward of the outer combustor wall; a primary zone, wherein within the primary zone the inner and outer combustor walls converge in a downstream direction; a secondary zone downstream of the primary zone, wherein in the secondary zone the inner and outer combustor walls are arranged to converge at a different rate to the primary zone, are non-convergent or are divergent in a downstream direction, such that a rate of change of radial width of the combustor is different in the primary zone to the secondary zone; a transition from the primary zone to the secondary zone; and a plurality of combustor cooling tiles lining the inner and outer combustor walls, wherein one or more of the tiles are arranged to extend from the primary zone to the secondary zone and across the transition from the primary zone to the secondary zone.

2 . The annular combustor according to claim 1 , wherein the tiles have a first portion provided in the primary zone, a second portion provided in the secondary zone and a transition region, the first portion, second portion and transition region being contiguous, and wherein the transition region of the tile has a greater radius of curvature than the inner and/or outer walls of the combustor at the transition between the primary and secondary zone.

3 . The annular combustor according to claim 1 , wherein a plurality of combustor cooling tiles are adjacently arranged in a circumferential direction to define an annulus that extends from the primary zone to the secondary zone across the transition from the primary zone to the secondary zone.

4 . The annular combustor according to claim 1 , wherein four tiles are adjacently arranged in a circumferential direction and together define a full annulus that extends from the primary zone to the secondary zone across the transition from the primary zone to the secondary zone of the inner and/or outer combustor wall.

5 . The annular combustor according to claim 1 , wherein the tile extending across the primary zone, secondary zone, and transition, extends across the entire primary zone and/or secondary zone.

6 . The annular combustor according to claim 1 , wherein the combustor cooling tiles are effusion tiles.

7 . The annular combustor according to claim 1 , wherein the inner and outer combustor walls are angled so as to reduce the radial width of the primary zone in a downstream direction.

8 . The annular combustor according to claim 1 , wherein the inner and outer combustor walls are angled so as to increase the radial width of the secondary zone in a downstream direction.

9 . A combustor for a gas turbine engine, the combustor comprising:
a primary zone; a secondary zone downstream of the primary zone; and a plurality of tiles lining the primary and secondary zone of the combustor, wherein a series of tiles are arranged to extend axially such that a portion of each of the tiles in said series of tiles is in the primary zone and a portion of said same tiles is in the secondary zone.

10 . The combustor according to claim 9 , wherein the combustor comprises an inner and outer combustor wall that converge at a first rate in the primary zone and are divergent, non-convergent or converge at a second rate in the secondary zone such that there is a change in axial direction of the walls at a transition between the primary and secondary zones, and wherein each of the tiles of the series of tiles are curved in a region coincident with the change in axial direction of the walls, the curve of the tiles having a greater radius than the change in direction of the walls.

11 . A combustor for a gas turbine engine, the combustor comprising:
inner and outer combustor walls defining a primary zone and a secondary zone, the inner and outer combustor walls in the primary zone being arranged at a first angle relative to each other, and the inner and outer combustor walls in the secondary zone being arranged at a second angle relative to each other, the second angle being different to the first angle; a fuel injector provided in the primary zone; an ignitor provided in the primary zone; a plurality of air inlets provided in the secondary zone for injecting air into the fit, combustor; and a plurality of combustor cooling tiles lining the combustor walls and connected thereto; wherein one or more of the combustor cooling tiles extend from the primary zone to the secondary zone such that a portion of the tile is in the primary zone, a portion of the tile is in the secondary zone and the tile extends across a transition from the primary zone to the secondary zone.

12 . A gas turbine engine comprising the combustor according to claim 1 .

13 . A gas turbine engine for an aircraft 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; 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; a combustor provided downstream of the compressor and upstream of the turbine, the combustor being the combustor according to claim 1 .

14 . The gas turbine engine according to claim 13 , 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.

15 . A combustor tile for a gas turbine engine, the combustor tile comprising:
a body having a first portion and a second portion, the first portion being contiguous with the second portion; and wherein the first portion is angled to the second portion by an angle between 185 and 210 degrees.

16 . The combustor tile according to claim 15 , wherein the first portion is angled to the second portion by an angle between 185 and 195 degrees.

17 . The combustor tile according to claim 15 , wherein a transition from the first portion to the second portion is curved such that there is a gradual change in angle of the tile from the first portion to the second portion.

Assignees

• Rolls Royce Plc

Inventors

• Hicks Robert

Classifications

• Y02T50/60

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

• F02K3/06

  with front fan · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F23R3/52 Primary

  Toroidal combustion chambers · CPC title

• F02C3/14

  characterised by the arrangement of the combustion chamber in the plant (combustion chambers per se F23R; F02C3/205 takes precedence) · CPC title