US3820324A - Flame tubes for gas turbine engines - Google Patents

Flame tubes for gas turbine engines Download PDF

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Publication number
US3820324A
US3820324A US00196160A US19616071A US3820324A US 3820324 A US3820324 A US 3820324A US 00196160 A US00196160 A US 00196160A US 19616071 A US19616071 A US 19616071A US 3820324 A US3820324 A US 3820324A
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US
United States
Prior art keywords
combustion chamber
openings
flame tube
wall
downstream
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.)
Expired - Lifetime
Application number
US00196160A
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English (en)
Inventor
W Grindley
G Bunn
A Ormerod
A Harrison
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
ZF International UK Ltd
Original Assignee
Lucas Industries Ltd
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
Application filed by Lucas Industries Ltd filed Critical Lucas Industries Ltd
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Publication of US3820324A publication Critical patent/US3820324A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/30—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices
    • F23R3/32—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices being tubular
    • 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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones

Definitions

  • a flame tube for a gas turbine engine includes a com bustion chamber having generally oppositely directed air openings which generate upstream and downstream vortex zones in the chamber.
  • Twin fuel sprayers supply fuel to the respective zones, whereby combustion conditions within each zone are separately maintained.
  • This invention relates to flame tubes for gas turbine engines and has as an object to provide a flame tube having a satisfactory combustion efficiency over a wide range of fuel flow rates.
  • a flame tube for a gas turbine engine comprises an annular casing, an inlet for the casing to which compressed air is, in use, supplied, an annular combustion chamber within the outer casing, a first set of openings in a wall of the combustion chamber adapted to direct an airflow into the combustion chamber generally radially thereof, a second set of openings in the chamber wall downstream of the said first set and oppositely directed thereto, an airflow through the said openings defining, in use, upstream and downstream toroidal vortex zones, and pairs of fuel supply means within the combustion chamber respectively operable to supply fuel to the upstream and downstream zones, the arrangement being such that the air/fuel ratio in each zone remains within acceptable limits, irrespective of the overall air/fuel ratio supplied to the engine.
  • FIG. 1 shows, somewhat diagrammatically, a part section through an annular flame tube
  • FIGS. 2, 5, 8 and are sections through alternative fuel supply means
  • FIGS. 3, 6, 9 and 11 are views on the correspondingly numbered arrows in FIGS. 2, 5, 8 and 10 respectively, and
  • FIGS. 4 and 7 are sections on the corresponding lines in FIGS. 2 and 5.
  • the flame tube shown in FIG. 1 has an annular outer casing 10 formed with an inlet 11 through which compressed air is, in use, supplied.
  • an annular combustion chamber 12 which is formed at its upstream end 13 so as to direct the greater part of the air into the annular spaces 14 between the chamber 12 and casing 10.
  • the upstream end 13 has an opening 15 via which a proportion of the air can enter a cavity 16 of the chamber 12.
  • Openings 18, 19 between the cavities l6, 17 respectively and the remainder of the chamber 12 have associated internal baffles 18a, 19a whereby air passing through the openings 18, 19 is directed along the inside wall of the chamber 12.
  • Additional openings 20 in the walls of the chamber 12 also include means for directing airflow along the inside wall of the chamber.
  • Openings 21 into the chamber 12 are directed generally radially of the chamber 12. Airflow through openings 21 combines with flow through openings 18, 19, 20 to create a pair of toroidal vortex zones 22, 23 which respectively define first and second combustion zones within the chamber 12. Pairs of fuel sprayers, one of which is shown generally at 24, are adapted to direct atomised fuel jets 25, 26 into the first and second comi bustion zones respectively.
  • the amount of fuel delivered by the jet 26 is increased with increasing power demanded of the engine, the maximum flow of the jet 26 being about three times that of the jet 25. At maximum fuel flow from both jets 25, 26 the air/fuel ratio in each zone does not fall below a level at which combustion is substantially complete.
  • the flame tube described thus permits a wide range of air/fuel ratios to be used to meet varying engine operating conditions, while maintaining a high combustion efficiency over the whole of this range. It has in practice been found that air/fuel ratios between and 42.4, respectively corresponding to stand-off and take-off conditions, may be used while maintaining combustion efficiency at higher than 99 percent.
  • the alternative fuel supply means shown in FIGS. 2 to 4 is a vapouriser arrangement 30 mounted in the wall of an annular flame tube 31.
  • Flame tube 31 is generally similar in form to the flame tube described with reference to FIG. 1 and operates as before to define a pair of toroidal vortex combustion zones 32, 33.
  • vapouriser arrangements 30 at angularly spaced positions around the flame tube 31.
  • the vapouriser arrangement 30 comprises a passage 34 and a further, crescent-section passage 35 having a common wall with, the passage 34.
  • Passage 34 has an outlet 37 extending through an opening 36 of the flame tube 31. Outlet 37 is directed radially of the flame tube 31 towards the combustion zone 33.
  • Passage 35 communicates with a transversely extending passage 38 which terminates in a pair of outlets 39, 40 directed towards the combustion zone 32.
  • Passages 34, 35 are formed so as to be substantially cylindrical externally and are surrounded by a sleeve 41 which provides a part annular air inlet 42. Outlets 37, 39, 40 and inlet 42 are positioned so that fluid flow therethrough will enhance the airflow within the flame tube creating the combustion zones 32, 33.
  • an air-fuel mixture is supplied to the passages 34, 35.
  • the fuel is vapourised by the heat of the flame tube and the air-vapour mixture is supplied to the combustion zones 32, 33.
  • vapouriser shown in FIGS. 5, 6 and 7 differs from that described above in that it is positioned centrally within a section of the annular flame tube. Passages 44, 45 have respective outlets 46, 47 positioned, as before, so that fluid flow therethrough reinforces the vortices in the respective combustion zones.
  • FIGS. 8 and 9 show yet another form of vapouriser.
  • Each of a plurality of vapourisers 50 extends radially into the annular flame tube 51 and comprises a tubular member 52, a pair of transverse tubes 53, 54 communieating with member 52 and a pair of outlets 55, 56 associated with the tubes 53, 54 respectively and directed radially outwards of the flame tube 51.
  • the member 52, tubes 53, 54 and outlets 55, 56 have a longitudinal division 57 to provide a pair of passages for fluid.
  • the vapourisers 50 are positioned within the flame tube 51 at a position intermediate the combustion zoneswhich are defined within the flame tube as previously described and fluid flowing through the outlets 55,56 acts, as before, to reinforce the existing vortices.
  • a fuel air mixture from the passages in the vapouriser 50 enters the adjacent combustion zone.
  • FIGS. and 11 Another means for supplying fuel to the combustion chamber of the invention is shown in FIGS. and 11.
  • An annular flame tube 60 is formed, as before, with air inlets which operate to define combustion zones 61, 62.
  • the flame tube 60 is also formed with an internal annular gutter 63 into which extends a plurality of fuel supply pipes 64.
  • the flame tube 60 is also formed with a plurality of air scoops 65 having radially directed downstream ends 66 and into which extend fuel pipes 67.
  • fuel enters the gutter 63 and scoops 65 and spills from the downstream openings thereof into the respective vortices of the zones 61, 62.
  • the fuel is at least partly vaporised by heat from the flame tube before spilling into the combustion zones.
  • a flame tube for a gas turbine engine comprising an annular outer casing, an inlet for the casing to which compressed air is, in use, supplied, an annular combustion chamber within the outer casing, said combustion chamber having an upstream end axially spaced from a downstream end thereof relative to the general direction of flow of gases through the flame tube in use, a first set of openings in a wall of the combustion chamber adapted to direct an air flow into the combustion chamber generally radially thereof, a second set of openings in the chamber wall downstream of the said first set and oppositely directed thereto, said first and second sets of openings being arranged so that, in use, an air flow therethrough finds a single upstream toroidal vortex zone and a single downstream toroidal vortex zone, and pairs of fuel supply means mounted in the wall of the combustion chamber and respectively operable to supply fuel to the upstream and downstream zones, the arrangement being such that the air/fuel ratio in each zone remains within acceptable limits, ir-
  • said flame tube including a third set of openings in the chamber wall and baffles within said wall adapted to direct air flow through said third openings along the inside of the wall, and further including means defining a cavity externally of said chamber and communicating therewith by means of said third openmgs.
  • a flame tube for a gas turbine engine comprising an annular outer casing, an inlet for the casing to which compressed air is, in use, supplied, an annular combustion chamber within the outer casing, said combustion chamberhaving an upstream end axially spaced from a downstream end thereof relative to the general direction of flow of gases through the flame tube in use, a first set of openings in a wall of the combustion chamber adapted to direct an air flow into the combustion chamber generally radially thereof, a second set of openings in the chamber wall downstream of the said first set and oppositely directed thereto, said first and second sets of openings being arranged so that, in use, an air flow therethrough finds a single upstream toroidal vortex zone and a single downstream toroidal vortex zone, and pairs of fuel supply means mounted in the wall of the combustion chamber and respectively operable to supply fuel to the upstream and downstream zones, wherein said fuel supply means comprises vapoun'ser arrangements, said flame tube including a third set of openings in the chamber wall and baffles within said wall
  • the flame tube according to claim 7 which includes a sleeve surrounding the said other passage and defining an air inlet for the combustion chamber.
  • a flame tube for a gas turbine engine comprising an annular outer casing, an inlet for the casing to which compressed air is, in use, supplied, an annular combustion chamber within the outer casing, said combustion chamber having an upstream end axially spaced from a downstream end thereof relative to the general direction of flow of gases through the flame tube in use, a first set of openings in a wall of the combustion chamber adapted to direct an air flow into the combustion chamber generally radially thereof, a second set of openings in the chamber wall downstream of the said first set and oppositely directed thereto, said first and second sets of openings being arranged so that, in use, an air flow therethrough finds a single upstream toroidal vortex zone and a single downstream toroidal vortex zone, and pairs of fuel supply means mounted in the wall of the combustion chamber and respectively operable to supply fuel to the upstream and downstream zones, said flame tube further including an annular gutter and a plurality of air scoops whose downstream ends are directed radially of the combustion chamber, and in which each pair of fuel supply

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Pressure-Spray And Ultrasonic-Wave- Spray Burners (AREA)
  • Spray-Type Burners (AREA)
  • Combustion Of Fluid Fuel (AREA)
US00196160A 1970-09-11 1971-11-05 Flame tubes for gas turbine engines Expired - Lifetime US3820324A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB4351070A GB1357533A (en) 1970-09-11 1970-09-11 Combustion equipment for gas turbine engines

Publications (1)

Publication Number Publication Date
US3820324A true US3820324A (en) 1974-06-28

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Application Number Title Priority Date Filing Date
US00196160A Expired - Lifetime US3820324A (en) 1970-09-11 1971-11-05 Flame tubes for gas turbine engines

Country Status (4)

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US (1) US3820324A (de)
DE (1) DE2157181C3 (de)
FR (1) FR2160272B1 (de)
GB (1) GB1357533A (de)

Cited By (19)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3973390A (en) * 1974-12-18 1976-08-10 United Technologies Corporation Combustor employing serially staged pilot combustion, fuel vaporization, and primary combustion zones
US4062182A (en) * 1974-12-21 1977-12-13 Mtu Motoren-Und Turbinen-Union Gmbh Combustion chamber for gas turbine engines
FR2363700A1 (fr) * 1976-09-04 1978-03-31 Rolls Royce Appareillage de combustion pour moteur a turbine a gaz
US4162611A (en) * 1976-07-07 1979-07-31 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Combustion chamber for turbo engines
US4179881A (en) * 1973-02-28 1979-12-25 United Technologies Corporation Premix combustor assembly
US4244179A (en) * 1977-01-28 1981-01-13 Kainov Gennady P Annular combustion chamber for gas turbine engines
US4271674A (en) * 1974-10-17 1981-06-09 United Technologies Corporation Premix combustor assembly
US4903492A (en) * 1988-09-07 1990-02-27 The United States Of America As Represented By The Secretary Of The Air Force Dilution air dispensing apparatus
US5265425A (en) * 1991-09-23 1993-11-30 General Electric Company Aero-slinger combustor
EP1524473A1 (de) * 2003-10-13 2005-04-20 Siemens Aktiengesellschaft Verfahren und Vorrichtung zum Verbrennen von Brennstoff
US20100212325A1 (en) * 2009-02-23 2010-08-26 Williams International, Co., L.L.C. Combustion system
US20120067051A1 (en) * 2009-12-29 2012-03-22 Oechsle Victor L Gas turbine engine and combustor
US20150040576A1 (en) * 2013-03-15 2015-02-12 Rolls-Royce Corporation Counter swirl doublet combustor
US20150159877A1 (en) * 2013-12-06 2015-06-11 General Electric Company Late lean injection manifold mixing system
CN105737204A (zh) * 2016-03-04 2016-07-06 武汉英康汇通电气有限公司 一种旋风火焰筒及其涡轮发电机
FR3066009A1 (fr) * 2017-05-02 2018-11-09 Safran Helicopter Engines Canne de prevaporisation pour une turbomachine
WO2020239702A1 (fr) * 2019-05-28 2020-12-03 Safran Helicopter Engines Canne de prévaporisation, ensemble de combustion muni de celle-ci et turbomachine munie de celui-ci
US20230194087A1 (en) * 2021-12-16 2023-06-22 General Electric Company Swirler opposed dilution with shaped and cooled fence
US20230204212A1 (en) * 2021-12-29 2023-06-29 Hanwha Aerospace Co., Ltd. Combustor

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
USRE30925E (en) 1977-12-14 1982-05-11 Caterpillar Tractor Co. Fuel vaporizing combustor tube
GB2102936B (en) * 1981-07-28 1985-02-13 Rolls Royce Fuel injector for gas turbine engines

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1130091A (fr) * 1954-05-06 1957-01-30 Nat Res Dev Perfectionnements apportés aux dispositifs de combustion
GB824306A (en) * 1956-04-25 1959-11-25 Rolls Royce Improvements in or relating to combustion equipment of gas-turbine engines
DE1097213B (de) * 1958-05-22 1961-01-12 Rolls Royce Verbrennungsanlage fuer Gasturbinentriebwerke
US3064424A (en) * 1959-09-30 1962-11-20 Gen Motors Corp Flame tube
FR1261034A (fr) * 1960-06-24 1961-05-12 Procédé et appareil pour le mélange de fluides
GB1136543A (en) * 1966-02-21 1968-12-11 Rolls Royce Liquid fuel combustion apparatus for gas turbine engines
FR1470246A (fr) * 1966-03-01 1967-02-17 Lucas Industries Ltd Moteur à turbine à gaz

Cited By (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4179881A (en) * 1973-02-28 1979-12-25 United Technologies Corporation Premix combustor assembly
US4271674A (en) * 1974-10-17 1981-06-09 United Technologies Corporation Premix combustor assembly
US3973390A (en) * 1974-12-18 1976-08-10 United Technologies Corporation Combustor employing serially staged pilot combustion, fuel vaporization, and primary combustion zones
US4062182A (en) * 1974-12-21 1977-12-13 Mtu Motoren-Und Turbinen-Union Gmbh Combustion chamber for gas turbine engines
US4162611A (en) * 1976-07-07 1979-07-31 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Combustion chamber for turbo engines
FR2363700A1 (fr) * 1976-09-04 1978-03-31 Rolls Royce Appareillage de combustion pour moteur a turbine a gaz
US4244179A (en) * 1977-01-28 1981-01-13 Kainov Gennady P Annular combustion chamber for gas turbine engines
US4903492A (en) * 1988-09-07 1990-02-27 The United States Of America As Represented By The Secretary Of The Air Force Dilution air dispensing apparatus
US5265425A (en) * 1991-09-23 1993-11-30 General Electric Company Aero-slinger combustor
EP1524473A1 (de) * 2003-10-13 2005-04-20 Siemens Aktiengesellschaft Verfahren und Vorrichtung zum Verbrennen von Brennstoff
WO2005038348A1 (de) * 2003-10-13 2005-04-28 Siemens Aktiengesellschaft Verfahren und vorrichtung zum verbrennen von brennstoff
US20070141519A1 (en) * 2003-10-13 2007-06-21 Siemens Aktiengesellschaft Method and device for the combustion of fuel
US20100212325A1 (en) * 2009-02-23 2010-08-26 Williams International, Co., L.L.C. Combustion system
US8640464B2 (en) * 2009-02-23 2014-02-04 Williams International Co., L.L.C. Combustion system
US9328924B2 (en) 2009-02-23 2016-05-03 Williams International Co., Llc Combustion system
US20120067051A1 (en) * 2009-12-29 2012-03-22 Oechsle Victor L Gas turbine engine and combustor
US8776525B2 (en) * 2009-12-29 2014-07-15 Rolls-Royce North American Technologies, Inc. Gas turbine engine and combustor
US20150040576A1 (en) * 2013-03-15 2015-02-12 Rolls-Royce Corporation Counter swirl doublet combustor
US9765969B2 (en) * 2013-03-15 2017-09-19 Rolls-Royce Corporation Counter swirl doublet combustor
US20150159877A1 (en) * 2013-12-06 2015-06-11 General Electric Company Late lean injection manifold mixing system
CN105737204A (zh) * 2016-03-04 2016-07-06 武汉英康汇通电气有限公司 一种旋风火焰筒及其涡轮发电机
CN105737204B (zh) * 2016-03-04 2018-01-26 武汉英康汇通电气有限公司 一种旋风火焰筒及其涡轮发电机
FR3066009A1 (fr) * 2017-05-02 2018-11-09 Safran Helicopter Engines Canne de prevaporisation pour une turbomachine
FR3096761A1 (fr) * 2019-05-28 2020-12-04 Safran Helicopter Engines Canne de prévaporisation, ensemble de combustion muni de celle-ci et turbomachine munie de celui-ci
WO2020239702A1 (fr) * 2019-05-28 2020-12-03 Safran Helicopter Engines Canne de prévaporisation, ensemble de combustion muni de celle-ci et turbomachine munie de celui-ci
CN113924445A (zh) * 2019-05-28 2022-01-11 赛峰直升机发动机公司 预蒸发管、设置有预蒸发管的燃烧组件以及设置有燃烧组件的涡轮机
US20220235937A1 (en) * 2019-05-28 2022-07-28 Safran Helicopter Engines Pre-vaporizing pipe, combustion assembly provided therewith and turbomachine provided therewith
CN113924445B (zh) * 2019-05-28 2023-11-21 赛峰直升机发动机公司 预蒸发管、设置有预蒸发管的燃烧组件以及设置有燃烧组件的涡轮机
US11867401B2 (en) * 2019-05-28 2024-01-09 Safran Helicopter Engines Pre-vaporizing pipe, combustion assembly provided therewith and turbomachine provided therewith
US20230194087A1 (en) * 2021-12-16 2023-06-22 General Electric Company Swirler opposed dilution with shaped and cooled fence
US11703225B2 (en) * 2021-12-16 2023-07-18 General Electric Company Swirler opposed dilution with shaped and cooled fence
US20230204212A1 (en) * 2021-12-29 2023-06-29 Hanwha Aerospace Co., Ltd. Combustor

Also Published As

Publication number Publication date
DE2157181A1 (de) 1973-05-24
GB1357533A (en) 1974-06-26
FR2160272A1 (de) 1973-06-29
DE2157181C3 (de) 1981-05-07
DE2157181B2 (de) 1980-09-18
FR2160272B1 (de) 1975-02-21

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