GB2444947A - Wall Element and Associated Structure for a Gas Turbine Engine - Google Patents
Wall Element and Associated Structure for a Gas Turbine Engine Download PDFInfo
- Publication number
- GB2444947A GB2444947A GB0625369A GB0625369A GB2444947A GB 2444947 A GB2444947 A GB 2444947A GB 0625369 A GB0625369 A GB 0625369A GB 0625369 A GB0625369 A GB 0625369A GB 2444947 A GB2444947 A GB 2444947A
- Authority
- GB
- United Kingdom
- Prior art keywords
- wall
- body portion
- duct
- combustor
- wall element
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000001816 cooling Methods 0.000 claims abstract description 18
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 18
- 230000008719 thickening Effects 0.000 claims abstract description 16
- 239000012530 fluid Substances 0.000 claims abstract description 7
- 230000004888 barrier function Effects 0.000 abstract 1
- 239000011248 coating agent Substances 0.000 abstract 1
- 238000000576 coating method Methods 0.000 abstract 1
- 238000002485 combustion reaction Methods 0.000 description 14
- 239000000446 fuel Substances 0.000 description 8
- 230000001141 propulsive effect Effects 0.000 description 3
- 239000000203 mixture Substances 0.000 description 2
- NJPPVKZQTLUDBO-UHFFFAOYSA-N novaluron Chemical compound C1=C(Cl)C(OC(F)(F)C(OC(F)(F)F)F)=CC=C1NC(=O)NC(=O)C1=C(F)C=CC=C1F NJPPVKZQTLUDBO-UHFFFAOYSA-N 0.000 description 2
- 238000013459 approach Methods 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- RLQJEEJISHYWON-UHFFFAOYSA-N flonicamid Chemical compound FC(F)(F)C1=CC=NC=C1C(=O)NCC#N RLQJEEJISHYWON-UHFFFAOYSA-N 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 230000000153 supplemental effect Effects 0.000 description 1
- 239000012720 thermal barrier coating Substances 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A wall structure for a gas turbine engine combustor has an outer wall 27 and an inner wall (28 fig 2) defining a duct between the inner and outer walls for the passage of cooling air, with the inner wall composed of several wall elements, in the form of tiles 29A and 29B. The outer wall has a radial step and the inner wall overlaps the radial step and has a local thickening region 50 positioned opposite the radial step. The inner wall element may have a body portion 36 aligned in use with the general direction of fluid flow through the combustor and a plurality of pedestals 38 that extend within the duct from the body portion towards the outer wall. The wall element may overlap adjacent elements at their upstream 30 and downstream 31 edge regions. Air may be admitted through the outer wall at an aperture 40, the air flow splitting into an upstream flow 42 and a downstream flow 44. The invention provides that the velocity of cooling air through the duct is maintained and cooling of the combustor wall elements is subsequently improved. The wall element may also include a barrier coating 64 which provides further heat resistance.
Description
* 2444947
WALL ELEMENTS FOR GAS TURBINE ENGINE COMPONENTS
This invention relates to wall elements for gas turbine engine combustors.
A typical gas turbine engine combustor includes a generally annular chamber having a plurality of fuel injectors at an upstream head end. Combustion air is provided through the head and through ports provided in the combustor walls downstream of the fuel injectors.
In order to improve the thrust and fuel consumption of gas turbine engines, i.e. the thermal efficiency, it is necessary to use high compressor pressures and combustion temperatures. Higher compressor pressures give rise to higher compressor outlet temperatures and higher pressures in the combustion chamber.
There is, therefore, a need to provide effective cooling of the combustion chamber walls. One cooling method which has been proposed is the provision of a double walled combustion chamber in which the inner wall is formed of a plurality of heat resistant tiles. Cooling air is directed into the duct between the outer walls and the tile from an aperture located midway along the tile. The flow of air bifurcates into upstream and downstream flows which are exhausted into the combustion chamber past the upstream and downstream edges of the tile. As the downstream flow approaches the end of the tile it is supplemented by air from the downstream tile before exiting to form a film over the downstream tile. The confluence of the flow with the flow from the downstream tile is typically at a region where the outer wall of the combustor steps radially. The radial step changes the velocity of the cooling air flow and affects the rate of heat removal at the rear edge of the tile which is also the location of the tile most susceptible to erosion.
According to the present invention there is provided a wall structure for an annular gas turbine engine corubustor arranged to have a general direction of fluid flow therethrough, the wall structure including an outer wall having a radial step and an inner wall overlapping the radial step, a duct being defined between the inner and outer walls for the passage of cooling air; the wall structure being characterised in that the inner wall has a local thickening opposing the radial step.
The outer wall may have a plurality of apertures for feeding cooling air into the duct.
Preferably the inner wall includes a plurality of wall elements, each wall element having a body portion aligned in use with the general direction of fluid flow through the combustor and a plurality of pedestals that extend within the duct from the body portion towards the outer wall.
Preferably the body portion provides the local thickening.
The downstream end of the body portion of an upstream wall element may overlap the upstream end of the body portion of a downstream wall element.
Preferably the local thickening has a contour that follows the contour of the radial step.
According to a second aspect of the invention there is provided a wall element for use as part of an inner wall of a gas turbine engine combustor wall structure including inner and outer walls, the inner and outer walls defining a duct therebetweeri, the wall element having a body portion aligned in use with a general direction of fluid flow through the combustor, the wall element having a local thickening in its downstream end region the local thickening being adapted to oppose a radial step in the outer wall.
Preferably the wall element has a plurality of pedestals arranged in use to extend within the duct from the body portion towards the outer wall.
Embodiments of the present invention will now be described by way of example only and with reference to the accompanying drawings, in which:-Fig. 1 is a sectional side view of the upper half of a gas turbine engine; Fig. 2 is a vertical cross-section through the combustor of the gas turbine engine shown in Fig. 1; Fig. 3 is a diagrammatic vertical cross-section through part of the wall structure of the combustor shown in Fig. 1.
Fig. 4 is a diagrammatic vertical cross-section of an alternative wall structure of the combustor shown in Fig. 1.
Fig. 5 is a diagrammatic vertical cross-section of an alternative wall structure of the combustor shown in Fig. 1.
Referring to Fig. 1, a gas turbine engine generally indicated at 10 has a principal axis X-X. The engine 10 comprises, in axial flow series, an air intake 11, d propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust nozzle 19.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produces two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbine 16, 17, 18 respectively drive the high and intermediate pressure compressors 14 and 13, and the fan 12 by suitable interconnecting shafts.
Referring to Fig. 2, the combustor 15 is constituted by an annular combustion chamber 20 having radially inner and outer wall structures 21 and 22 respectively. The combustion chamber 20 is secured to an engine casing 23 by a plurality of pins 24 (only one of which is shown). Fuel is directed into the chamber 20 through a number of injector nozzles 25 (only one of which is shown) located at the upstream end of the combustion chamber 20. Fuel injector nozzles 25 are circumferentially spaced around the engine 10 and serve to spray fuel into air delivered from the high pressure compressor 1.4. The resulting fuel/air mixture is then combusted within the chamber 20.
The combustion process which takes place generates a large amount of heat. It is therefore necessary to arrange that the inner and outer wall structures 21 and 22 are capable of withstanding this heat.
The inner and outer wall structures 21 and 22 are generally of the same construction and comprise an outer wall 27 and an inner wall 28. The inner wall 28 is made up of a plurality of discrete wall elements in the form of tiles 29, which are all of the same general rectangular configuration and are positioned adjacent each other. The circumferentially extending edges 30,31 of adjacent tiles overlap each other. Each tile 29 is provided with threaded studs 32 which project through apertures in the outer wall 27. Nuts 34 are screwed onto threaded studs 32 and tightened against the outer wall 27, thereby securing the tiles 29 in place.
Both the radially outer and inner outer walls 27 of the annular cornbustor have a series of radial steps that enable optimum use of the cooling air. Air which has passed through the pedestals of an upstream tile is relatively cool and can be used for film cooling the downstream combustor, which must be offset to present the file face at the exit point of the air flow emanating from the upst'ream tile. The step additionally strengthens the combustor against buckling under flame out or surge. The upstream end of a tile 29 lies adjacent the step whilst the downstream end of the upstream tile axially overlaps both the radial step and the upstream end of the downstream tile.
Referring to Fig. 3, there is shown part of the inner wall structure 21 showing two overlapping tiles. 29A, 29B.
Each of the tiles 29A, 29B comprises a main body portion 36 which, in combination with the main body portions of each of the other tiles 22, defines the inner wall 28. A plurality of heat removal members in the form of upstanding substantially cylindrical pedestals 38 extend from each body member 36 towards the inner wall of the combustor 27 - which forms the outer wall of the combustor wall sructure.
The downstream edge region 31 of tile 29A overlaps the upstream edge region 30 of tile 29B.
The body member 36 and outer wall of the wall structure 27 define a duct 37 that extends therebetween.
Cooling air is supplied to the duct 37 through an aperture extending through the outer wall 27. The flow bifurcates to provide an upstream flow 42 that flows substantially in the opposite direction to the general flow of combustion gasses through the combustor and a downstream flow 44 that flows generally in the same direction as combustion gasses through the combustor.
The body member has a thermal barrier coating 64 on the surface facing the combustion chamber 20 to provide further heat resistance.
At the downstream end of tile 29A the downstream flow mixes with the upstream flow from tile 29B and is then exhausted as a film of cooling air over the combustor facing surface of the body member 36 of tile 29B. The confluence of the flows occurs where the outer wall 27 of the combustor wall structure steps radially.
To avoid an excessive reduction in the velocity of the air flow through the duct at this point the body member 36 has a circumferentially arranged local thickening 50 which follows the radial step of the outer cold-skin wall 27. The thickening is contoured to a maxima before reducing as it extends axially rearward. This enables a relatively constant velocity across the whole length of the duct 37 thereby maintaining a relatively high heat removal rate, which drops if the velocity of cooling air flow drops significantly. The high heat removal is therefore maintained particularly at the downstream edge region of a tile where the tile temperature peaks and the tile integrity is at greatest risk.
Pedestals 38 are provided on the region of local thickening 50. The length of the pedestals is maintained over the hump shaped local thickening maintaining the high heat removal afforded by these structures. The pedestals define a flow-path for the supplemental air from the downstream tile and which maximjses the volume flow of cooling air within the pedestal array.
Various modifications may be made without departing from the scope of the invention. For example, the degree of axial overlap of the upstream and downstream tiles may be varied to optimise the film of air over the downstream tile. Similarly, the pedestal length in the region of the hump could be adjusted to optimise heat removal and the shape of the hump / local thickening could be refined to maintain the optimum cooling air velocity.
Claims (9)
1. A wall structure for an annular gas turbine engine combustor arranged to have a general direction of fluid flow therethrough, the wall structure including an outer wall having a radial step and an inner wall overlapping the radial step, a duct being defined between the inner and outer walls for the passage of cooling air; the wall structure being characterised in that the inner wall has a local thickening opposing the radial step.
2. A wall structure according to claim 1, wherein the outer wall has a plurality of apertures for feeding cooling air into the duct.
3. A wall structure according to claim 1 or claim 2, wherein the inner wall includes a plurality of wall elements, each wall element having a body portion aligned in use with the general direction of fluid flow through the combustor and a plurality of pedestals that extend within the duct from the body portion towards the outer wall.
4. A wall structure according to claim 3, wherein the body portion provides the local thickening.
5. A wall structure according to claim 3, or claim 4, wherein the downstream end of the body portion of an upstream wall element overlaps the upstream end of the body portion of a downstream wall element.
6. A wall structure according to any preceding claim, wherein the local thickening has a contour that follows the contour of the radial step.
7. A wall element for use as part of an inner wall of a gas turbine engine combustor wall structure including inner and outer walls, the inner and outer walls defining a duct therebetween, the wall element having a body portion aligned in use with a general direction of fluid flow through the combustor, the wall element having a local thickening in its downstream end region the local thickening being adapted to oppose a radial step in the outer wall.
8. A wall element according to claim 7, wherein the wall element has a plurality of pedestals arranged in use to extend within the duct from the body portion towards the outer wall.
9. A wall structure and / or wall element as hereinbe ore described with reference to the accompanying drawings.
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB0625369A GB2444947B (en) | 2006-12-19 | 2006-12-19 | Wall elements for gas turbine engine components |
US11/987,892 US20080145211A1 (en) | 2006-12-19 | 2007-12-05 | Wall elements for gas turbine engine components |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB0625369A GB2444947B (en) | 2006-12-19 | 2006-12-19 | Wall elements for gas turbine engine components |
Publications (3)
Publication Number | Publication Date |
---|---|
GB0625369D0 GB0625369D0 (en) | 2007-01-31 |
GB2444947A true GB2444947A (en) | 2008-06-25 |
GB2444947B GB2444947B (en) | 2009-04-08 |
Family
ID=37734496
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
GB0625369A Expired - Fee Related GB2444947B (en) | 2006-12-19 | 2006-12-19 | Wall elements for gas turbine engine components |
Country Status (2)
Country | Link |
---|---|
US (1) | US20080145211A1 (en) |
GB (1) | GB2444947B (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20150362192A1 (en) * | 2013-01-17 | 2015-12-17 | United Technologies Corporation | Gas turbine engine combustor liner assembly with convergent hyperbolic profile |
Families Citing this family (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8813718B2 (en) | 2008-12-31 | 2014-08-26 | Speed Of Air, Inc. | Internal combustion engine |
GB201114745D0 (en) * | 2011-08-26 | 2011-10-12 | Rolls Royce Plc | Wall elements for gas turbine engines |
US20160238249A1 (en) * | 2013-10-18 | 2016-08-18 | United Technologies Corporation | Combustor wall having cooling element(s) within a cooling cavity |
US10767863B2 (en) | 2015-07-22 | 2020-09-08 | Rolls-Royce North American Technologies, Inc. | Combustor tile with monolithic inserts |
US10386067B2 (en) * | 2016-09-15 | 2019-08-20 | United Technologies Corporation | Wall panel assembly for a gas turbine engine |
US11125434B2 (en) * | 2018-12-10 | 2021-09-21 | Raytheon Technologies Corporation | Preferential flow distribution for gas turbine engine component |
Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2179276A (en) * | 1983-12-19 | 1987-03-04 | Gen Electric | Fabricated metal panel and method |
GB2355301A (en) * | 1999-10-13 | 2001-04-18 | Rolls Royce Plc | A wall structure for a combustor of a gas turbine engine |
Family Cites Families (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4653279A (en) * | 1985-01-07 | 1987-03-31 | United Technologies Corporation | Integral refilmer lip for floatwall panels |
FR2714152B1 (en) * | 1993-12-22 | 1996-01-19 | Snecma | Device for fixing a thermal protection tile in a combustion chamber. |
FR2752916B1 (en) * | 1996-09-05 | 1998-10-02 | Snecma | THERMAL PROTECTIVE SHIRT FOR TURBOREACTOR COMBUSTION CHAMBER |
GB9803291D0 (en) * | 1998-02-18 | 1998-04-08 | Chapman H C | Combustion apparatus |
US6438958B1 (en) * | 2000-02-28 | 2002-08-27 | General Electric Company | Apparatus for reducing heat load in combustor panels |
GB2359882B (en) * | 2000-02-29 | 2004-01-07 | Rolls Royce Plc | Wall elements for gas turbine engine combustors |
GB2368902A (en) * | 2000-11-11 | 2002-05-15 | Rolls Royce Plc | A double wall combustor arrangement |
ES2309029T3 (en) * | 2001-01-09 | 2008-12-16 | Mitsubishi Heavy Industries, Ltd. | GAS TURBINE COMBUSTION CHAMBER. |
GB2384046B (en) * | 2002-01-15 | 2005-07-06 | Rolls Royce Plc | A double wall combuster tile arrangement |
EP1482246A1 (en) * | 2003-05-30 | 2004-12-01 | Siemens Aktiengesellschaft | Combustion chamber |
-
2006
- 2006-12-19 GB GB0625369A patent/GB2444947B/en not_active Expired - Fee Related
-
2007
- 2007-12-05 US US11/987,892 patent/US20080145211A1/en not_active Abandoned
Patent Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2179276A (en) * | 1983-12-19 | 1987-03-04 | Gen Electric | Fabricated metal panel and method |
GB2355301A (en) * | 1999-10-13 | 2001-04-18 | Rolls Royce Plc | A wall structure for a combustor of a gas turbine engine |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20150362192A1 (en) * | 2013-01-17 | 2015-12-17 | United Technologies Corporation | Gas turbine engine combustor liner assembly with convergent hyperbolic profile |
Also Published As
Publication number | Publication date |
---|---|
GB2444947B (en) | 2009-04-08 |
GB0625369D0 (en) | 2007-01-31 |
US20080145211A1 (en) | 2008-06-19 |
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Legal Events
Date | Code | Title | Description |
---|---|---|---|
PCNP | Patent ceased through non-payment of renewal fee |
Effective date: 20201219 |