WO1992016798A1 - Gas turbine engine combustor - Google Patents

Gas turbine engine combustor Download PDF

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Publication number
WO1992016798A1
WO1992016798A1 PCT/GB1992/000201 GB9200201W WO9216798A1 WO 1992016798 A1 WO1992016798 A1 WO 1992016798A1 GB 9200201 W GB9200201 W GB 9200201W WO 9216798 A1 WO9216798 A1 WO 9216798A1
Authority
WO
WIPO (PCT)
Prior art keywords
wall
gas turbine
turbine engine
wall elements
elements
Prior art date
Application number
PCT/GB1992/000201
Other languages
French (fr)
Inventor
Anthony Pidcock
Stephen Mark Cooper
Peter Fry
Original Assignee
Rolls-Royce Plc
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Family has litigation
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Application filed by Rolls-Royce Plc filed Critical Rolls-Royce Plc
Priority to JP4504028A priority Critical patent/JPH06507468A/en
Priority to EP92904247A priority patent/EP0576435B1/en
Priority to DE69204280T priority patent/DE69204280T2/en
Publication of WO1992016798A1 publication Critical patent/WO1992016798A1/en

Links

Classifications

    • 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
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2260/00—Function
    • F05B2260/20—Heat transfer, e.g. cooling
    • F05B2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2260/00—Function
    • F05B2260/20—Heat transfer, e.g. cooling
    • F05B2260/202—Heat transfer, e.g. cooling by film 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

  • GAS TURBINE ENGINE COMBUSTOR This invention relates to a gas turbine engine combustor and in particular to the construction of the wall of such a combustor.
  • the wall is made up of two parts: a continuous outer wall and an inner wall made up of a number of partially overlapping inner wall elements.
  • the outer wall and inner wall elements are maintained in spaced apart relationship and cooling air is directed through holes in the outer wall into the space defined between them.
  • the cooling air flows through the space to be exhausted through gaps defined between the overlapping portions of the inner wall elements.
  • the cooling air thereby provides convection cooling as it flows between the inner wall elements and outer wall and film cooling of the inner wall elements after it has been exhausted from the gaps between inner wall elements.
  • a gas turbine engine combustor is provided with a wall structure which comprises an outer wall and an inner wall, said inner wall being constituted by a plurality of discreet wall elements, means being provided to maintain said wall elements and said outer wall in spaced apart relationship, said outer wall being apertured to permit the flow of cooling fluid into the spaced defined between said outer wall and said wall elements, each of said wall elements being apertured to facilitate the exhaustion of said cooling fluid from said spaces, means being provided to interconnect the periphery of each wall element and said outer wall said interconnection means defining a continuous wall around each wall element periphery so that a discreet chamber is thereby defined between each of said wall elements and said outer wall for the flow therethrough of said cooling fluid.
  • Figure 1 is a sectional side view of the upper half of a ducted fan gas turbine engine which incorporates a combustor in accordance with the present invention
  • Figure 2 is a sectional side view of a portion of the wall of the combustor of the gas turbine engine shown in figure 1;
  • Figure 3 is a view on arrow A of figure 2;
  • Figure 4 is a view on an enlarged scale of a portion of the combustor wall shown in figure 2;
  • Figure 5 is a view on arrow B of figure 4.
  • Figure 6 is a view similar to Figure 2 showing a modified form of combustor in accordance with the present invention.
  • a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 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 the conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second airflow which provides propulsive thrust.
  • the intermediate pressure compressor 13 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 combustion equipment 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 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust.
  • the high, intermediate and low pressure turbines 16,17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.
  • the combustion equipment 15 is constituted by an annular combustor 20 having radially inner and outer wall structures 21 and 22 respectively. Fuel is directed into the combustor 20 through a number of fuel nozzles (not shown) located at the upstream end 23 of the combustor 20. The fuel nozzles are circumferentially spaced around the engine 10 and serve to spray fuel into air derived from the high pressure compressor 14. The resultant fuel/air mixture is them combusted within the combustor 20.
  • the radially outer wall structure 22 can be seen more clearly if reference is now made to figure 2. It will be appreciated, however, that the radially inner wall structure 21 is of the same general configuration as the radially outer wall structure 22.
  • the radially outer wall structure 22 comprises an outer wall 24 and an inner wall 25.
  • the inner wall 25 is made up of a plurality of discreet wall elements 26 which are all of the same general rectangular configuration and are positioned adjacent each other. The majority of each wall element 26 is arranged to 5 be equi-distant from the outer wall 24. . However, the periphery of each wall element 26 is provided with a continuous flange 27 to facilitate the spacing apart of the wall element 26 and the outer wall 24. It will be seen therefore that a chamber 28 is thereby defined between each 10 wall element 26 and the outer wall 24.
  • Each wall element 26 is of cast construction and is provided with integral bolts 29 which facilitate its attachment to the outer wall 24.
  • That air is then exhausted from the chambers 28 through a plurality of angled effusion holes 32 provided in each wall element 26.
  • the effusion holes 32 are are so angled as to be aligned in a generally downstream C direction with regard to the general fluid flow through the combustor 20.
  • each of the wall elements 26 is provided with two highly effective forms of cooling: impingement cooling and film cooling. They are therefore fully protected from the effects of the high temperatures within the combustor 20.
  • a further feature of the present invention is that none of the wall elements 26 presents exposed edges to the combustion process within the combustor 20. Consequently the overheating problems which may be experienced with wall elements having such exposed edges are avoided.
  • pedestals 33 are integral with the wall elements 26 and engage or terminate very close to the outer wall 24.
  • the provision of the pedestals 33, which tend to be located in the central region of each wall element 26, results in a reduction in the number of the angled effusion holes 32 in each wall element 26. Consequently, the angled effusion holes 32 tend to be concentrated in the edge regions of the wall elements 26.

Abstract

A gas turbine engine combustor (20) is provided with a wall structure (22) which comprises an outer wall (24) having a plurality of wall elements (26) attached thereto. Each wall element (26) is provided with a flange (27) around its periphery which serves to define a chamber (28) between each wall element (26) and the outer wall (24). Holes (30) in the outer wall (24) permit the flow of cooling air into each chamber (28) to provide impingement cooling of the wall elements (26). Holes (30) in the wall elements (26) permit the exhaustion of cooling air from the chambers (28) to provide film cooling of the wall elements (26).

Description

GAS TURBINE ENGINE COMBUSTOR This invention relates to a gas turbine engine combustor and in particular to the construction of the wall of such a combustor.
The combustion1 process which takes place within the combustor of a gas turbine engine results in the combustor walls being exposed to extremely high temperatures. The alloys used in combustor wall construction are normally unable to withstand these temperatures without some form of cooling. Various combustor wall designs have been employed in the past which make use of pressurised air derived from the engine compressor for cooling purposes. In one particular wall design described in Great Britain Patent
Application No 2,087,065A, the wall is made up of two parts: a continuous outer wall and an inner wall made up of a number of partially overlapping inner wall elements.
The outer wall and inner wall elements are maintained in spaced apart relationship and cooling air is directed through holes in the outer wall into the space defined between them. The cooling air flows through the space to be exhausted through gaps defined between the overlapping portions of the inner wall elements. The cooling air thereby provides convection cooling as it flows between the inner wall elements and outer wall and film cooling of the inner wall elements after it has been exhausted from the gaps between inner wall elements.
It has been found with combustion chamber walls of this type that the film cooling of the inner wall elements is not as effective as would normally be desired. This can lead to overheating of and possible damage to the exposed edges of the overlapping portions of the inner wall elements.
It is an object of the present invention to provide a gas turbine engine combustor wall construction in which such film cooling is of improved effectiveness.
According to the present invention, a gas turbine engine combustor is provided with a wall structure which comprises an outer wall and an inner wall, said inner wall being constituted by a plurality of discreet wall elements, means being provided to maintain said wall elements and said outer wall in spaced apart relationship, said outer wall being apertured to permit the flow of cooling fluid into the spaced defined between said outer wall and said wall elements, each of said wall elements being apertured to facilitate the exhaustion of said cooling fluid from said spaces, means being provided to interconnect the periphery of each wall element and said outer wall said interconnection means defining a continuous wall around each wall element periphery so that a discreet chamber is thereby defined between each of said wall elements and said outer wall for the flow therethrough of said cooling fluid.
The present invention will now be described, by way of example, with reference to the accompanying drawings:
Figure 1 is a sectional side view of the upper half of a ducted fan gas turbine engine which incorporates a combustor in accordance with the present invention;
Figure 2 is a sectional side view of a portion of the wall of the combustor of the gas turbine engine shown in figure 1;
Figure 3 is a view on arrow A of figure 2; Figure 4 is a view on an enlarged scale of a portion of the combustor wall shown in figure 2; Figure 5 is a view on arrow B of figure 4.
Figure 6 is a view similar to Figure 2 showing a modified form of combustor in accordance with the present invention.
With reference to figure 1 a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 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 the conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second airflow which provides propulsive thrust. The intermediate pressure compressor 13 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 combustion equipment 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 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16,17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.
The combustion equipment 15 is constituted by an annular combustor 20 having radially inner and outer wall structures 21 and 22 respectively. Fuel is directed into the combustor 20 through a number of fuel nozzles (not shown) located at the upstream end 23 of the combustor 20. The fuel nozzles are circumferentially spaced around the engine 10 and serve to spray fuel into air derived from the high pressure compressor 14. The resultant fuel/air mixture is them combusted within the combustor 20.
The combustion process which takes place within the combustion 20 naturally generates a large amount of heat. It is necessary therefore to arrange that the inner and outer wall structures 21 and 22 are capable of withstanding this heat while functioning in a normal manner.
The radially outer wall structure 22 can be seen more clearly if reference is now made to figure 2. It will be appreciated, however, that the radially inner wall structure 21 is of the same general configuration as the radially outer wall structure 22.
Referring to figure 2, the radially outer wall structure 22 comprises an outer wall 24 and an inner wall 25. The inner wall 25 is made up of a plurality of discreet wall elements 26 which are all of the same general rectangular configuration and are positioned adjacent each other. The majority of each wall element 26 is arranged to 5 be equi-distant from the outer wall 24. . However, the periphery of each wall element 26 is provided with a continuous flange 27 to facilitate the spacing apart of the wall element 26 and the outer wall 24. It will be seen therefore that a chamber 28 is thereby defined between each 10 wall element 26 and the outer wall 24.
Each wall element 26 is of cast construction and is provided with integral bolts 29 which facilitate its attachment to the outer wall 24.
During engine operation, some of the air exhausted from 5 the high pressure compressor 14 is permitted to flow over the exterior -surfaces of the combustor 20. The air provides combustor 20 cooling and some of it is directed into the interior of the combustor 20 to assist in the combustion process. A large number of holes 30 are 0 provided in the outer wall 24, which can also be seen in figure 3, to permit the flow of some of this air into the chambers 28. The air passing through the holes 30 impinges upon the radially outward surfaces of the wall elements 26 as indicated by the air flow indicating arrows 31. This 5 ensures that each of the wall elements 26 is cooled in a highly effective manner. That air is then exhausted from the chambers 28 through a plurality of angled effusion holes 32 provided in each wall element 26. The effusion holes 32 are are so angled as to be aligned in a generally downstream C direction with regard to the general fluid flow through the combustor 20.
The angled effusion holes 32, which can be seen more clearly in figures 4 and 5, are not of circular cross-sectional shape. Instead they are all of the = so-called race-track configuration, that is, they have two parallel sides interconnected by semi-circular cross-section portions. This shape, together with the inclination of the hole 32, ensures that air exhausted from them forms a film of cooling air over the inward surface of each wall element 26, that is, the surface which confronts the combustion process which takes place within the combustor 20. This film of cooling air assists in protecting the wall elements 26 from the effects of the high temperature gases within the combustor 20.
It will be appreciated that although the present invention has been described with reference to effusion holes 32 which are of race-track cross-sectional configuration, other alternative configurations may also be effective in providing satisfactory wall element 26 cooling.
It will be seen therefore that each of the wall elements 26 is provided with two highly effective forms of cooling: impingement cooling and film cooling. They are therefore fully protected from the effects of the high temperatures within the combustor 20.
A further feature of the present invention is that none of the wall elements 26 presents exposed edges to the combustion process within the combustor 20. Consequently the overheating problems which may be experienced with wall elements having such exposed edges are avoided.
It may be desirable in certain circumstances to enhance the heat exchange relationship between the cooling air passing through the chambers 28 and the wall elements 26. One way of readily achieving this would be to provide pedestals 33 or other suitable devices to increase surface area on the surfaces of the wall elements 26 which confront the outer wall 24 as can be seen in Figure 6. The pedestals 33 are integral with the wall elements 26 and engage or terminate very close to the outer wall 24. The provision of the pedestals 33, which tend to be located in the central region of each wall element 26, results in a reduction in the number of the angled effusion holes 32 in each wall element 26. Consequently, the angled effusion holes 32 tend to be concentrated in the edge regions of the wall elements 26.

Claims

Claims:-
1. A gas turbine engine combustor (15) provided with a wall structure (22) which comprises an outer wall (24) and an inner wall (25), said inner wall (25) being constituted by a plurality of discreet wall elements (26), means being provided to maintain the majority of said wall elements (26) and said outer wall (24) in spaced apart relationship, said outer wall (24) being apertured to permit the flow of cooling fluid into the spaces defined between said outer wall (24) and said wall elements (26), each of said wall elements (26) being apertured to facilitate the exhaustion of said cooling fluid from said spaces, characterised in that means (27) are provided to interconnect the periphery of each wall element (26) and said outer wall (24), said interconnection means (27) defining a continuous wall around each wall element (26) periphery so that a discreet chamber (28) is defined between each of said wall elements (26) and said outer wall (24) for the flow therethrough of said cooling fluid.
2. A gas turbine engine combustor as claimed in claim 1 characterised in that said apertures (30) in said outer wall
(24) are so arranged as to direct cooling fluid on to said wall elements (26) to provide impingement cooling thereof.
3. A gas turbine engine combustor as claimed in claim 1 or claim 2 characterised in that said apertures (32) in each of said wall elements (25) are so arranged as to exhaust cooling fluid from said discreet chambers (28) to provide film cooling of said wall elements (25).
4. A gas turbine engine combustor as claimed in claim 3 characterised in that said apertures (32) in said wall elements (25) are inclined in the general direction of fluid flow through said combustor to facilitate said film cooling of said wall elements (25).
5. A gas turbine engine combustor as claimed in claim 4 characterised in that said apertures (32) in said wall elements (25) are of race-track cross-sectional configuration.
6. A gas turbine engine combustor as claimed in any one preceding claim characterised in that said wall elements (25) are positioned on said outer wall (24) so as to be generally adjacent each other.
7. A gas turbine 'engine combustor as claimed in any one preceding claim characterised in that said means (27) to interconnect the periphery of each wall element (25) and said outer wall (24) comprises a continuous flange extending along the whole of the periphery of each of said wall elements (25).
8. A gas turbine engine combustor as claimed in claim 7 characterised in that the peripheral flange (27) on each of said vt&kl elements (25) additionally constitutes said means to space apart its associated wall element (25) and said outer wall (24).
9. A gas turbine engine combustor as claimed in any one preceding claim characterised in that each of said wall elements (25) is provided with integral bolts (29) to facilitate its attachment to said outer wall (24).
10. A gas turbine engine combustor as claimed in any one preceding claim characterised in that each of said wall elements (25) is provided with a plurality of pedestals (33) to enhance the heat exchange relationship between said wall elements (25) and said cooling fluid flow through said spaces (28) between said wall elements (25) and said outer wall (24).
11. A gas turbine engine combustor as claimed in claim 10 characterised in that each of said pedestals (33) engages said outer wall (24).
12. A gas turbine engine combustor as claimed in any one preceding claim characterised in that said combustor (15) is annular.
PCT/GB1992/000201 1991-03-22 1992-02-03 Gas turbine engine combustor WO1992016798A1 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
JP4504028A JPH06507468A (en) 1991-03-22 1992-02-03 gas turbine engine combustor
EP92904247A EP0576435B1 (en) 1991-03-22 1992-02-03 Gas turbine engine combustor
DE69204280T DE69204280T2 (en) 1991-03-22 1992-02-03 GAS TURBINE COMBUSTION CHAMBER.

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB9106085.5 1991-03-22
GB919106085A GB9106085D0 (en) 1991-03-22 1991-03-22 Gas turbine engine combustor

Publications (1)

Publication Number Publication Date
WO1992016798A1 true WO1992016798A1 (en) 1992-10-01

Family

ID=10692006

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/GB1992/000201 WO1992016798A1 (en) 1991-03-22 1992-02-03 Gas turbine engine combustor

Country Status (5)

Country Link
EP (1) EP0576435B1 (en)
JP (1) JPH06507468A (en)
DE (1) DE69204280T2 (en)
GB (1) GB9106085D0 (en)
WO (1) WO1992016798A1 (en)

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0694739A1 (en) * 1994-07-27 1996-01-31 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Double-walled combustor
GB2298266A (en) * 1995-02-23 1996-08-28 Rolls Royce Plc A cooling arrangement for heat resistant tiles in a gas turbine engine combustor
EP0741268A1 (en) * 1995-05-03 1996-11-06 United Technologies Corporation Liner panel for a gas turbine combustor wall
DE19516798A1 (en) * 1995-05-08 1996-11-14 Abb Management Ag Premix burner with axial or radial air flow
FR2752916A1 (en) * 1996-09-05 1998-03-06 Snecma THERMAL PROTECTIVE SHIRT FOR TURBOREACTOR COMBUSTION CHAMBER
WO1999063274A1 (en) * 1998-06-03 1999-12-09 Pratt & Whitney Canada Corp. Impingement and film cooling for gas turbine combustor walls
DE102007018061A1 (en) 2007-04-17 2008-10-23 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber wall
CN101922354A (en) * 2009-04-16 2010-12-22 通用电气公司 Turbogenerator with lining
EP2273196A2 (en) 2009-07-08 2011-01-12 Rolls-Royce Deutschland Ltd & Co KG Gas turbine combustion chamber head
CN102562309A (en) * 2010-12-21 2012-07-11 株式会社东芝 Transition piece and gas turbine
CN102607028A (en) * 2011-01-14 2012-07-25 通用电气公司 Apparatus for vibration support in combustors and method for forming apparatus
EP2559942A1 (en) 2011-08-19 2013-02-20 Rolls-Royce Deutschland Ltd & Co KG Gas turbine combustion chamber head with cooling and damping
EP2700877A2 (en) 2012-08-21 2014-02-26 Rolls-Royce Deutschland Ltd & Co KG Gas turbine combustion chamber with impingement-cooled bolts for the combustion chamber shingles
EP2749816A2 (en) 2012-12-27 2014-07-02 Rolls-Royce Deutschland Ltd & Co KG Method for arranging of impingement cooling holes and effusion holes in a combustion chamber wall of a gas turbine
EP2770260A2 (en) 2013-02-26 2014-08-27 Rolls-Royce Deutschland Ltd & Co KG Impact effusion cooled shingle of a gas turbine combustion chamber with elongated effusion bore holes
US9422830B2 (en) 2013-12-18 2016-08-23 Rolls-Royce Deutschland Ltd & Co Kg Washer of a combustion chamber tile of a gas turbine
US10551067B2 (en) 2011-11-10 2020-02-04 Ihi Corporation Combustor liner with dual wall cooling structure

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CH703657A1 (en) * 2010-08-27 2012-02-29 Alstom Technology Ltd Method for operating a burner arrangement and burner arrangement for implementing the process.
GB201105790D0 (en) 2011-04-06 2011-05-18 Rolls Royce Plc A cooled double walled article
DE102013222932A1 (en) 2013-11-11 2015-05-28 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with shingle for carrying out a spark plug
DE102016222099A1 (en) 2016-11-10 2018-05-17 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber of a gas turbine

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GB2204672A (en) * 1987-05-06 1988-11-16 Rolls Royce Plc Combustor
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GB1093515A (en) * 1966-04-06 1967-12-06 Rolls Royce Method of producing combustion chambers and similar components for gas turbine engines
US3422620A (en) * 1967-05-04 1969-01-21 Westinghouse Electric Corp Combustion apparatus
FR2333126A1 (en) * 1975-11-29 1977-06-24 Rolls Royce GAS TURBINE ENGINE COMBUSTION CHAMBER REFRIGERATION DEVICE
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Cited By (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0694739A1 (en) * 1994-07-27 1996-01-31 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Double-walled combustor
FR2723177A1 (en) * 1994-07-27 1996-02-02 Snecma Sa COMBUSTION CHAMBER COMPRISING A DOUBLE WALL
US5598697A (en) * 1994-07-27 1997-02-04 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Double wall construction for a gas turbine combustion chamber
GB2298266A (en) * 1995-02-23 1996-08-28 Rolls Royce Plc A cooling arrangement for heat resistant tiles in a gas turbine engine combustor
EP0741268A1 (en) * 1995-05-03 1996-11-06 United Technologies Corporation Liner panel for a gas turbine combustor wall
DE19516798A1 (en) * 1995-05-08 1996-11-14 Abb Management Ag Premix burner with axial or radial air flow
US5738509A (en) * 1995-05-08 1998-04-14 Asea Brown Boveri Ag Premix burner having axial or radial air inflow
FR2752916A1 (en) * 1996-09-05 1998-03-06 Snecma THERMAL PROTECTIVE SHIRT FOR TURBOREACTOR COMBUSTION CHAMBER
US6029455A (en) * 1996-09-05 2000-02-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Turbojet engine combustion chamber with heat protecting lining
WO1999063274A1 (en) * 1998-06-03 1999-12-09 Pratt & Whitney Canada Corp. Impingement and film cooling for gas turbine combustor walls
US8099961B2 (en) 2007-04-17 2012-01-24 Rolls-Royce Deutschland Ltd & Co Kg Gas-turbine combustion chamber wall
DE102007018061A1 (en) 2007-04-17 2008-10-23 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber wall
CN101922354A (en) * 2009-04-16 2010-12-22 通用电气公司 Turbogenerator with lining
US8677757B2 (en) 2009-07-08 2014-03-25 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber head of a gas turbine
EP2273196A2 (en) 2009-07-08 2011-01-12 Rolls-Royce Deutschland Ltd & Co KG Gas turbine combustion chamber head
DE102009032277A1 (en) 2009-07-08 2011-01-20 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber head of a gas turbine
CN102562309A (en) * 2010-12-21 2012-07-11 株式会社东芝 Transition piece and gas turbine
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EP2749816A2 (en) 2012-12-27 2014-07-02 Rolls-Royce Deutschland Ltd & Co KG Method for arranging of impingement cooling holes and effusion holes in a combustion chamber wall of a gas turbine
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Also Published As

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EP0576435A1 (en) 1994-01-05
GB9106085D0 (en) 1991-05-08
DE69204280T2 (en) 1996-01-25
JPH06507468A (en) 1994-08-25
DE69204280D1 (en) 1995-09-28
EP0576435B1 (en) 1995-08-23

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