US3691766A - Combustion chambers - Google Patents

Combustion chambers Download PDF

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Publication number
US3691766A
US3691766A US98837A US3691766DA US3691766A US 3691766 A US3691766 A US 3691766A US 98837 A US98837 A US 98837A US 3691766D A US3691766D A US 3691766DA US 3691766 A US3691766 A US 3691766A
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Prior art keywords
combustion chamber
passage
igniter
combustion
head
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US98837A
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Keith Harold Champion
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Rolls Royce PLC
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Rolls Royce PLC
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/26—Starting; Ignition
    • F02C7/264—Ignition
    • 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
    • F23R3/54—Reverse-flow combustion chambers

Definitions

  • the invention pertains to combustion apparatus for a gas turbine engine comprising a combustion chamber lying within a casing defining therewith a passage for the supply of airto the combustion chamber, an igniter situated in a recess formed in the wall of the easing, and a baffle situated in the passage adjacent and upstream of the igniter so that air flowing through the passage to the combustion chamber is induced by the baffle to form a stabilized vortex around the igniter.
  • This invention relates to combustion apparatus gas turbine engines.
  • combustion apparatus comprising an air casing connected to receive air from the compressor and a combustion chamber situated within the air casing and apertured to receive air therefrom.
  • the combustion chamber exhausts through the turbine.
  • a gas turbine engine combustion apparatus comprising a first wall structure defining a combustion chamber, a second wall structure defining an air casing surrounding the combustion chamber, a passage defined between said wall structures and through which air flows in a determined direction, a baffle arranged in the passage transversely to said direction for inducing vortices at the downstream side of the baffle, and a torch igniter connected to the casing in a position downstream of the baffle for presenting fuel and ignition to said vortices for generating a pilot flame.
  • the combustion chamber may be formed to define a head so related to the igniter that at least a part of the air flow past the igniter passes from one side of the head to the other on its way to one or more apertures in the chamber.
  • the casing wall structure is formed to define a recess open to the passage and located directly downstream of the baffle, the igniter being situated within said recess.
  • FIG. 1 is part sectional side elevation of a gas turbine engine.
  • FIG. 2 is a sectional view of the reverse flow combustion chamber shown in FIG. 1, drawn to an enlarged scale.
  • FIG. 3 is a part view in the direction of arrow C in FIG. 2.
  • a gas turbine engine 11 comprising a centrifugal compressor 12, combustion apparatus 13 and a turbine 14 all in flow series as shown.
  • the combustion apparatus 13 comprises an annular combustion chamber 15 mounted within an annular casing 16 and supported from an anchor point 17 on the mainframe 18 of the engine.
  • a main fuel injector 19 comprise a Hike pipe' 20 connected to the wall of the combustion chamber 15 and projecting thereinto.
  • the injector 19 includes a fuel supply pipe 21 connected to a fuel supply system 34 and passing through a hole 22 in the casing 16 and then through a hole 23 in the combustion chamber and terminating within the pipe 20.
  • a flexible metallic seal 24 prevents air leaks between thepipe 21 and the hole 22.
  • Air supplied by the compressor 12 flows, as indicated by arrows 25, along a passage 26, defined between the combustion chamber 15 and the casing 16, into the combustion chamber through ports 30 and then exits through an outlet 27 of the combustion chamber to the turbine 14.
  • a torch igniter 28 comprising an igniter plug 31 and fuel injector 29 connected to an ignition supply 35 and an auxiliary fuel injector 31 connected to the fuel supply system 34, is situated in a recess 32 in the casing 16.
  • a baffle 33 is attached to-a wall of the recess at a point upstream from the torch igniter and projects into the passage 26,
  • the air flowing from the compressor through the passage 25 is induced by the baffle to form stable vortices 36 around the torch igniter to create a suitable environment for ignition to be effected and the resulting flame at the auxiliary fuel injector to be maintained.
  • This pilot flame is continuously fed by the supply 34 and travels with the air along the passage 26 and into the combustion chamber 15, through ports 30 therein, where it ignites the mixture of air and the fuel from the main fuel injector 19.
  • the location of the igniter 28 is in a relatively narrow part of the passage 26 and upstream of the main fuel injector 19.
  • the narrowness of the passage in this example about 0.3 inch, requires that the baffle 33 is substantially in the form of a flat plate inducing a pair of generally twodimensional vortices 36, and said narrowness therefore also requires the presence of the recess 32 to provide for an area of slow air movement necessary, together with the vortices, to maintain burning.
  • the igniter 28 Since the igniter 28 is situated upstream of the injector 19 the pilot flame passes over the main fuel supply pipe 21 so as to raise the fuel temperature and improve vaporization in the pipe 20.
  • the igniter 28 is of course only actuated for starting of the engine, and is switched off once the main combustion is established.
  • pilot flame is distributed by the flow 25 around the head of the chamber, i.e. the area surrounding the injector 19 with generally beneficial results as regards raising the temperature of that area during starting.
  • the combustion apparatus is of the reverse flow type, that is, the mean direction of flow 36 within the chamber 15 is opposite to the general direction of flow 37 through the engine.
  • the position of the igniter 28 must be upstream of the primary combustion zone, i.e. the zone in the chamber 15 surrounding the main fuel injector 19 and constituting the head, denoted 39, of the chamber. This means that the igniter must be situated in the relatively narrow passage 26 and the baffle and recess make this possible as already mentioned.
  • this location of the igniter in a reverse flow chamber makes it possible for the pilot flame not only to travel over the main fuel injector 19 but also through the bend of the passage 26 around the primary zone and to the opposite side of the combustion chamber before entering the latter. In other words, the arrangement provides access by the pilot flame to either side of the head of the combustion chamber.
  • Combustion apparatus for a gas turbine engine comprising a first wall structure defining a combustion chamber, a second wall structure defining an air casing surrounding the combustion chamber, a passage defined between said wall structures and through which air flows in a predetermined direction, a recess formed in said second wall structure and opens towards said passage, a baffle arranged in-said passage at the upstream end of said recess and transversely to said direction of air flow for inducing vortices at the downstream side of said baffle, and a torch ignitersituated within said recess for presenting fuel and ignition to said vortices for generating a pilot flame.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)

Abstract

The invention pertains to combustion apparatus for a gas turbine engine comprising a combustion chamber lying within a casing defining therewith a passage for the supply of air to the combustion chamber, an igniter situated in a recess formed in the wall of the casing, and a baffle situated in the passage adjacent and upstream of the igniter so that air flowing through the passage to the combustion chamber is induced by the baffle to form a stabilized vortex around the igniter.

Description

United States Patent Champion [54] COMBUSTION CHAMBERS [72] Inventor: Keith Harold Champion, Rickmansworth, England I [73] Assignee: Rolls-Royce Limited, Derby, En-
gland 22 Filed: Dec. 16, 1970 21 Appl. No.: 98,837
[52] US. Cl. .Q ..60/39.82 P, 60/39.65 [51] Int. Cl. ..F02c 7/26 [58] Field of Search.....60/39.82 P, 39.82 S, 39.82 R,
[56] References Cited UNlTED STATES PATENTS 2,715,816 8/1955 Thorn et a1 ..60/39.65 2,592,110 4/1952 Berggren et a1. ..60/39.82 P 2,621,477 12/1952 Powter et al. ..60/39.82 P 3,540,216 11/1970 Quillevere et al. ....60/39.82 P
1451 Sept. 19, 1972 3,124,933 3/1964 Stram et al. ..60/39.82 P
FOREIGN PATENTS OR APPLlCATlONS 1,476,843 6/1969 Germany ..60/39.82 P 644,719 10/ 1950 Great Britain ..60/39.23 201,132 l/l955 Austria ..60/39.82 R
Primary Examiner-Carlton R. Croyle Assistant Examiner-Warren Olsen Attorney-Stevens, Davis, Miller & Mosher 5 7] ABSTRACT The invention pertains to combustion apparatus for a gas turbine engine comprising a combustion chamber lying within a casing defining therewith a passage for the supply of airto the combustion chamber, an igniter situated in a recess formed in the wall of the easing, and a baffle situated in the passage adjacent and upstream of the igniter so that air flowing through the passage to the combustion chamber is induced by the baffle to form a stabilized vortex around the igniter.
' 2 Claims, 3 Drawing Figures COMBUSTION CHAMBERS This invention relates to combustion apparatus gas turbine engines.
It is known in gas turbine engines to provide a compressor, combustion apparatus and turbine in flow series, the combustion apparatus comprising an air casing connected to receive air from the compressor and a combustion chamber situated within the air casing and apertured to receive air therefrom. The combustion chamber exhausts through the turbine.
It is known to provide a torch igniter in the space between the air casing and combustion chamber so as to form a pilot flame which is carried by the air flow into the chamber there to ignite fuel supplied by a main fuel supply.
In relatively small engines where the passage formed between the casing and the combustion chamber is a narrow one, for example less than 0.5 inch, there are difficulties in establishing stable combustion conditions for the torch igniter. It is an object of the invention to provide a combustion apparatus in which these combustion conditions in a relatively narrow space can be successfully maintained. It is also an object of the invention to provide an improved distribution of the pilot flame in said passage with a view to aiding the main combustion process.
According to this invention there is provided a gas turbine engine combustion apparatus comprising a first wall structure defining a combustion chamber, a second wall structure defining an air casing surrounding the combustion chamber, a passage defined between said wall structures and through which air flows in a determined direction, a baffle arranged in the passage transversely to said direction for inducing vortices at the downstream side of the baffle, and a torch igniter connected to the casing in a position downstream of the baffle for presenting fuel and ignition to said vortices for generating a pilot flame.
The combustion chamber may be formed to define a head so related to the igniter that at least a part of the air flow past the igniter passes from one side of the head to the other on its way to one or more apertures in the chamber.
Preferably the casing wall structure is formed to define a recess open to the passage and located directly downstream of the baffle, the igniter being situated within said recess.
An embodiment of the invention will now be described with reference to the accompanying drawings in which:
FIG. 1 is part sectional side elevation of a gas turbine engine.
FIG. 2 is a sectional view of the reverse flow combustion chamber shown in FIG. 1, drawn to an enlarged scale.
FIG. 3 is a part view in the direction of arrow C in FIG. 2.
Referring to the drawings, there is shown a gas turbine engine 11 comprising a centrifugal compressor 12, combustion apparatus 13 and a turbine 14 all in flow series as shown.
The combustion apparatus 13 comprises an annular combustion chamber 15 mounted within an annular casing 16 and supported from an anchor point 17 on the mainframe 18 of the engine.
A main fuel injector 19 comprise a Hike pipe' 20 connected to the wall of the combustion chamber 15 and projecting thereinto. The injector 19 includes a fuel supply pipe 21 connected to a fuel supply system 34 and passing through a hole 22 in the casing 16 and then through a hole 23 in the combustion chamber and terminating within the pipe 20. A flexible metallic seal 24 prevents air leaks between thepipe 21 and the hole 22.
Air supplied by the compressor 12 flows, as indicated by arrows 25, along a passage 26, defined between the combustion chamber 15 and the casing 16, into the combustion chamber through ports 30 and then exits through an outlet 27 of the combustion chamber to the turbine 14.
A torch igniter 28, comprising an igniter plug 31 and fuel injector 29 connected to an ignition supply 35 and an auxiliary fuel injector 31 connected to the fuel supply system 34, is situated in a recess 32 in the casing 16. A baffle 33 is attached to-a wall of the recess at a point upstream from the torch igniter and projects into the passage 26,
In operation the air flowing from the compressor through the passage 25 is induced by the baffle to form stable vortices 36 around the torch igniter to create a suitable environment for ignition to be effected and the resulting flame at the auxiliary fuel injector to be maintained. This pilot flame is continuously fed by the supply 34 and travels with the air along the passage 26 and into the combustion chamber 15, through ports 30 therein, where it ignites the mixture of air and the fuel from the main fuel injector 19.
The location of the igniter 28 is in a relatively narrow part of the passage 26 and upstream of the main fuel injector 19.
The narrowness of the passage, in this example about 0.3 inch, requires that the baffle 33 is substantially in the form of a flat plate inducing a pair of generally twodimensional vortices 36, and said narrowness therefore also requires the presence of the recess 32 to provide for an area of slow air movement necessary, together with the vortices, to maintain burning.
Since the igniter 28 is situated upstream of the injector 19 the pilot flame passes over the main fuel supply pipe 21 so as to raise the fuel temperature and improve vaporization in the pipe 20.
The igniter 28 is of course only actuated for starting of the engine, and is switched off once the main combustion is established.
Further, it will be seen that the pilot flame is distributed by the flow 25 around the head of the chamber, i.e. the area surrounding the injector 19 with generally beneficial results as regards raising the temperature of that area during starting.
The combustion apparatus is of the reverse flow type, that is, the mean direction of flow 36 within the chamber 15 is opposite to the general direction of flow 37 through the engine. The position of the igniter 28 must be upstream of the primary combustion zone, i.e. the zone in the chamber 15 surrounding the main fuel injector 19 and constituting the head, denoted 39, of the chamber. This means that the igniter must be situated in the relatively narrow passage 26 and the baffle and recess make this possible as already mentioned. At the same time this location of the igniter in a reverse flow chamber makes it possible for the pilot flame not only to travel over the main fuel injector 19 but also through the bend of the passage 26 around the primary zone and to the opposite side of the combustion chamber before entering the latter. In other words, the arrangement provides access by the pilot flame to either side of the head of the combustion chamber.
We claim:
1. Combustion apparatus for a gas turbine engine comprising a first wall structure defining a combustion chamber, a second wall structure defining an air casing surrounding the combustion chamber, a passage defined between said wall structures and through which air flows in a predetermined direction, a recess formed in said second wall structure and opens towards said passage, a baffle arranged in-said passage at the upstream end of said recess and transversely to said direction of air flow for inducing vortices at the downstream side of said baffle, and a torch ignitersituated within said recess for presenting fuel and ignition to said vortices for generating a pilot flame.
2. Apparatus according to claim 1 wherein said combustion chamber has a head defining a primary c0mbustion zone and extending therefrom so that the mean direction of flow in the combustion chamber is opposite to said predetermined direction, the passage extending around said head so that the pilot flame can pass from one side of the head adjacent the igniter to the opposite side of the head.

Claims (2)

1. Combustion apparatus for a gas turbine engine comprising a first wall structure defining a combustion chamber, a second wall strucTure defining an air casing surrounding the combustion chamber, a passage defined between said wall structures and through which air flows in a predetermined direction, a recess formed in said second wall structure and opens towards said passage, a baffle arranged in said passage at the upstream end of said recess and transversely to said direction of air flow for inducing vortices at the downstream side of said baffle, and a torch igniter situated within said recess for presenting fuel and ignition to said vortices for generating a pilot flame.
2. Apparatus according to claim 1 wherein said combustion chamber has a head defining a primary combustion zone and extending therefrom so that the mean direction of flow in the combustion chamber is opposite to said predetermined direction, the passage extending around said head so that the pilot flame can pass from one side of the head adjacent the igniter to the opposite side of the head.
US98837A 1970-12-16 1970-12-16 Combustion chambers Expired - Lifetime US3691766A (en)

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Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2410737A1 (en) * 1977-12-01 1979-06-29 United Technologies Corp BURNER FOR GAS TURBINE
US4192139A (en) * 1976-07-02 1980-03-11 Volkswagenwerk Aktiengesellschaft Combustion chamber for gas turbines
US4549402A (en) * 1982-05-26 1985-10-29 Pratt & Whitney Aircraft Of Canada Limited Combustor for a gas turbine engine
WO1989006308A1 (en) * 1987-12-28 1989-07-13 Sundstrand Corporation Annular combustor with tangential cooling air injection
US5085039A (en) * 1989-12-07 1992-02-04 Sundstrand Corporation Coanda phenomena combustor for a turbine engine
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
EP0539580A1 (en) * 1991-05-13 1993-05-05 Sundstrand Corp Very high altitude turbine combustor.
US5237813A (en) * 1992-08-21 1993-08-24 Allied-Signal Inc. Annular combustor with outer transition liner cooling
US5628193A (en) * 1994-09-16 1997-05-13 Alliedsignal Inc. Combustor-to-turbine transition assembly
US6055813A (en) * 1997-08-30 2000-05-02 Asea Brown Boveri Ag Plenum
US6269628B1 (en) * 1999-06-10 2001-08-07 Pratt & Whitney Canada Corp. Apparatus for reducing combustor exit duct cooling
EP1160432A1 (en) * 2000-05-31 2001-12-05 Daniel Bregentzer Gas turbine engine
US20040088988A1 (en) * 2002-11-08 2004-05-13 Swaffar R. Glenn Gas turbine engine transition liner assembly and repair
US20060101828A1 (en) * 2004-11-16 2006-05-18 Patel Bhawan B Low cost gas turbine combustor construction
EP1847779A2 (en) * 2006-04-21 2007-10-24 Honeywell International Inc. Optimized configuration of a reverse flow combustion system for a gas turbine engine
US20120328996A1 (en) * 2011-06-23 2012-12-27 United Technologies Corporation Reverse Flow Combustor Duct Attachment
US20130008168A1 (en) * 2010-03-26 2013-01-10 Matthias Hase Burner for stabilizing the combustion of a gas turbine
US20210102704A1 (en) * 2019-10-04 2021-04-08 United Technologies Corporation Engine turbine support structure
US11415058B2 (en) * 2020-12-23 2022-08-16 Collins Engine Nozzles, Inc. Torch ignitors with tangential injection
US11415059B2 (en) * 2020-12-23 2022-08-16 Collins Engine Nozzles, Inc. Tangentially mounted torch ignitors
US20220412561A1 (en) * 2021-06-28 2022-12-29 Delavan Inc. Passive secondary air assist nozzles

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4192139A (en) * 1976-07-02 1980-03-11 Volkswagenwerk Aktiengesellschaft Combustion chamber for gas turbines
US4168609A (en) * 1977-12-01 1979-09-25 United Technologies Corporation Folded-over pilot burner
FR2410737A1 (en) * 1977-12-01 1979-06-29 United Technologies Corp BURNER FOR GAS TURBINE
US4549402A (en) * 1982-05-26 1985-10-29 Pratt & Whitney Aircraft Of Canada Limited Combustor for a gas turbine engine
USRE34962E (en) * 1987-12-28 1995-06-13 Sundstrand Corporation Annular combustor with tangential cooling air injection
WO1989006308A1 (en) * 1987-12-28 1989-07-13 Sundstrand Corporation Annular combustor with tangential cooling air injection
US4928479A (en) * 1987-12-28 1990-05-29 Sundstrand Corporation Annular combustor with tangential cooling air injection
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5085039A (en) * 1989-12-07 1992-02-04 Sundstrand Corporation Coanda phenomena combustor for a turbine engine
EP0539580A1 (en) * 1991-05-13 1993-05-05 Sundstrand Corp Very high altitude turbine combustor.
EP0539580A4 (en) * 1991-05-13 1993-12-15 Sundstrand Corporation, Inc. Very high altitude turbine combustor
US5237813A (en) * 1992-08-21 1993-08-24 Allied-Signal Inc. Annular combustor with outer transition liner cooling
US5628193A (en) * 1994-09-16 1997-05-13 Alliedsignal Inc. Combustor-to-turbine transition assembly
US6055813A (en) * 1997-08-30 2000-05-02 Asea Brown Boveri Ag Plenum
US6269628B1 (en) * 1999-06-10 2001-08-07 Pratt & Whitney Canada Corp. Apparatus for reducing combustor exit duct cooling
JP2003502546A (en) * 1999-06-10 2003-01-21 プラット アンド ホイットニー カナダ コーポレイション Combustor outlet duct cooling reduction device
EP1160432A1 (en) * 2000-05-31 2001-12-05 Daniel Bregentzer Gas turbine engine
US6553765B2 (en) 2000-05-31 2003-04-29 Daniel Bregentzer Turbojet engine
US20040088988A1 (en) * 2002-11-08 2004-05-13 Swaffar R. Glenn Gas turbine engine transition liner assembly and repair
US6925810B2 (en) * 2002-11-08 2005-08-09 Honeywell International, Inc. Gas turbine engine transition liner assembly and repair
US20060101828A1 (en) * 2004-11-16 2006-05-18 Patel Bhawan B Low cost gas turbine combustor construction
US7350358B2 (en) * 2004-11-16 2008-04-01 Pratt & Whitney Canada Corp. Exit duct of annular reverse flow combustor and method of making the same
EP1847779A2 (en) * 2006-04-21 2007-10-24 Honeywell International Inc. Optimized configuration of a reverse flow combustion system for a gas turbine engine
US20070245710A1 (en) * 2006-04-21 2007-10-25 Honeywell International, Inc. Optimized configuration of a reverse flow combustion system for a gas turbine engine
EP1847779A3 (en) * 2006-04-21 2008-08-13 Honeywell International Inc. Optimized configuration of a reverse flow combustion system for a gas turbine engine
US20130008168A1 (en) * 2010-03-26 2013-01-10 Matthias Hase Burner for stabilizing the combustion of a gas turbine
US8864492B2 (en) * 2011-06-23 2014-10-21 United Technologies Corporation Reverse flow combustor duct attachment
US20120328996A1 (en) * 2011-06-23 2012-12-27 United Technologies Corporation Reverse Flow Combustor Duct Attachment
US20210102704A1 (en) * 2019-10-04 2021-04-08 United Technologies Corporation Engine turbine support structure
US11753952B2 (en) * 2019-10-04 2023-09-12 Raytheon Technologies Corporation Support structure for a turbine vane of a gas turbine engine
US11415058B2 (en) * 2020-12-23 2022-08-16 Collins Engine Nozzles, Inc. Torch ignitors with tangential injection
US11415059B2 (en) * 2020-12-23 2022-08-16 Collins Engine Nozzles, Inc. Tangentially mounted torch ignitors
US20220412561A1 (en) * 2021-06-28 2022-12-29 Delavan Inc. Passive secondary air assist nozzles
US11543130B1 (en) * 2021-06-28 2023-01-03 Collins Engine Nozzles, Inc. Passive secondary air assist nozzles
US20230097301A1 (en) * 2021-06-28 2023-03-30 Collins Engine Nozzles, Inc. Passive secondary air assist nozzles
US11859821B2 (en) * 2021-06-28 2024-01-02 Collins Engine Nozzles, Inc. Passive secondary air assist nozzles

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