US3826082A - Combustion liner cooling slot stabilizing dimple - Google Patents

Combustion liner cooling slot stabilizing dimple Download PDF

Info

Publication number
US3826082A
US3826082A US00346595A US34659573A US3826082A US 3826082 A US3826082 A US 3826082A US 00346595 A US00346595 A US 00346595A US 34659573 A US34659573 A US 34659573A US 3826082 A US3826082 A US 3826082A
Authority
US
United States
Prior art keywords
liner
lip
downstream
improvement
space
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
US00346595A
Other languages
English (en)
Inventor
R Smuland
R Ward
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Priority to US00346595A priority Critical patent/US3826082A/en
Priority to CA192,405A priority patent/CA996761A/en
Priority to DE2414376A priority patent/DE2414376A1/de
Priority to IT49727/74A priority patent/IT1003927B/it
Priority to BE142678A priority patent/BE813091A/xx
Priority to JP49034778A priority patent/JPS5025918A/ja
Priority to FR7411194A priority patent/FR2223554A1/fr
Application granted granted Critical
Publication of US3826082A publication Critical patent/US3826082A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • F23R3/08—Arrangement of apertures along the flame tube between annular flame tube sections, e.g. flame tubes with telescopic sections

Definitions

  • ABSTRACT A combustor-liner film cooling slot, of the variety including an annular, space extending circumferentially' of the liner which space is defined between overlapping portions of telescoping liner segments, a cooling fluid plenum substantially circumscribing the liner, means for transferring cooling fluid from the plenum to the space, means for exhausting fluid from the space onto the liner, and an overhanging lip extending substantially axially downstream of the space for facilitating attachment of the fluid to the liner in a protective film'barrier, is provided with a plurality of circumferentiallyspaced stabilizing dimples disposed within the overhanging lip.
  • Each dimple includes a radially depressed portion which has'a larger circumferential width at its upstream extremity than its downstream extremity.
  • the depression is disposed toward the associated cooler liner segment and is provided with a geometry which enhances stress relief of the lip as well as cooling fluid attachment as a protective film barrier upon the heated side of the liner downstream from the lip.
  • This invention relates to combustion chambersand, more particularly, to means for effecitively cooling combustion chambers.
  • the presentinvention will be discussed in the environment of combustion chambers from a plurality of individual jets, corresponding to the 5 individual apertures, into a flow having substantially uniform circumferential velocity.
  • a combustion chamber liner defining a combustion zone also partially defines a cool fluid plenum usually circumscribing the combustion zone. Means are commonly provided for transferring a portion of the cool fluid from the plenum into the combustion zone to form the protective film barrier from the hot combustion products in the chamber.
  • the present invention provides an improvement.
  • the present invention in one embodiment thereof, provides a combustion chamber liner which comprises a number of telescoping, partially overlapping segments.
  • a cooling fluid plenum substantially circumscribes the liner, and means are provided for transferring cooling fluid from the plenum to an annular space defined between adjacent overlapping liner seg ments.
  • the radially inner of the overlapping segments I forms an overhanging lip extending substantially axially downstream of the space which facilitates attachment of the fluid to the liner in a protective film barrier.
  • This lip is provided with a plurality of circumferentially spaced stabilizing dimples, each dimple including a radially depressed portion which has an upstream, substantially linear extremity (or junction with the lip surface) and a downstream arcuate extremity, which occurs at the downstream edge of the lip.
  • Each dimple has a larger width in the circumferential direction at its upstream extremity than at its downstream extremity.
  • each dimple departs from the lip surface to a first depth near its upstream extremity and to a second depth near its downstream extremity, the second depth being greater than the first.
  • Each dimple also includes lateral extremities at the junction of the depressed portion with the lip surface, the lateral extremities of any one dimple converging in the downstream direction. As a result, the adjacent of the lateral extremities of adjacent dimples diverge in the downstream direction.
  • the depression extends in the direction of the associated liner segment.
  • FIG. 1 is a simplified cross-sectional view of a cornbustion chamber of a gas turbine engine employing cooling slots
  • FIG. 2 depicts'a cross-sectional view of areinforcing dimple according to the prior art
  • FIG. 3 illustrates a pictorial view of the prior art dimple of FIG. 2 taken along line 3-3;
  • FIG. 4 shows a cross-sectional view of a cooling slot according to the present invention
  • FIG. 5 is a pictorial view taken along line 4-4'of FIG. 3 and showing details of the reinforcing dimple of the present invention
  • FIG. 6 is a graphical representation of stress levels in prior art and dimples according to the present invention.
  • FIG. 7 is a graphical representation of comparative aerodynamic efficiencies.
  • the combustion zone itself is designated 44 and is defined by liners 34 and 40 as well as by an upstream dome 46 which cooperates with a fuel nozzle 48 through fuel for combustion is directed into the combustion zone.
  • An air/fuel inlet 50 is defined between axial extensions 52 and 54 of liners 34 and 40, respectively.
  • combustion chamber described is substantially similar to those in present use.
  • a flow of atmospheric air is pressurized by means of a compressor (not shown) upstream of the combustion zone 44 with the compressor discharge directed partially into plena 52 and 54 as well as into the fuel/air inlet 50.
  • a quantity of fuel is mixed with the portion of air entering fuel/air inlet 50 and is ignited within combustion zone 44.
  • the rapid expansion of the burning gases and the configuration of liners 34 and 40 results in the. gases being forced from combustion zone 44 serves to drive the upstream compressor through an interconnecting shaft.
  • the remaining energy of the gas stream provides energy for driving thrust toward the left in FIG. 1.
  • FIG. 2 illustrates a crosssectional view of a cooling slot according to the prior art.
  • FIG. 3 is a view of this same slot taken along line 3-3 of FIG. 2 and further particularizes the cooling slot of the prior art.
  • liner 34 may be seen to be divided substantially into a number of axially adjacent segments. A typical segment 60 can beseen to be in telescopic cooperation with a typical segment 62? downstream thereof by means of a junction designated generally by the numeral 64'.
  • a cooling slot configuration which, in substance, comprises a cooling film promoter for passing cool fluid from plenum 36' in a protective film barrier upon liner 34, the latter liner partially defining a hot gas passage (the combustion zone 44).
  • the segment 60 can be seen to partially define the plenum 36 as well as the. combustion zone 44'.
  • Segment 62' likewise partially defines the plenum and the hot gas passage, and these liner segments cooperate to form an annular space 66' extending circumferentially of the liner between them.
  • the space 66' includes asubstantially closed upstream end 70f and a downstream-facing open end or exit 68'. Thus, the space is substantially isolated from the combustion zone except for communication through exit 68.
  • the annular space 66 is disposed between a portion of downstream segment 62' and an xial extension of the upstream segment 60' of liner 34'. It can further be seen that the space 66' is connected to the surrounding plenum 36 by means of a plurality of circumferentially spaced apertures 74' through a portion of the liner. It is the function of these apertures to deliver cool fluid from the plenum to the space.
  • the barrier formationto provide a more uniform means for delivering cool fluid to the space than the plurality of apertures 74'.
  • the cool fluid quantity must be metered, and these apertures are a most effective means for accomplishing metering.
  • the structural forces imposed upon the liner require that it be formed in a particularly strong fashion.
  • a more uniform delivery system such as a continuous annular slot between the plenum 36 and space 66 would substantially weaken the liner in comparison to apertures 74.'I-Ience, the apertures are a preferred delivery mechanism.
  • the apertures do serve to transfer the fluid to the space 66 in a plurality of spaced jets with subthrough an outlet 56 and into engagement with a turbine 58.
  • Rotary portions of the turbine are driven by this exiting fluid and -a portion of the energy thereof stantial circumferential velocity gradients therebetween.
  • the overhanging lip 72 extends a substantial distance downstream of the junction 76' between liner segments 60' and 62.
  • the space 66' is given a substantial axial length.
  • the fluid passing from plenum 36' in a plurality of jets through apertures 74' is given a predetermined period of residence within space 66' as it passes downstream therein.
  • the viscous forces within the fluid and frictional engagement thereof with the surfaces of the lip 72' and downstream liner segment 62' cause the fluid to develop a more uniform circumferential velocity.
  • the space 66 must be substantial in length and, hence, the overhanging lip 72 must likewise extend axially downstream for a substantial distance. In this way, the overhanging lip 72 serves to facilitate the attachment of the cool fluid to the liner in a protective film barrier.
  • the depressed portion 92 of the dimple extends radially from the surface of lip' 72 in the direction outwardly of the center of the combustion zone and toward the associated downstream liner segment 62.
  • the depressed portion departs from the surface of the lip to a first depth near its upstream extremity and to a second and greater depth near its downstream extremity. In other words, the depressed portion of the dimple converges toward the downstream liner segment 62 in the downstream direction.
  • the axial length of the lip is such that-residualstresses within the structure defining the lip combine withtherr'nal stresses resulting from the heat of combustion within chamber 44' to cause structural deformation of the lip that can result in local closing of the exit 68', through which the cool fluid is normally exhausted from the space onto the liner. Should this occur, localized hot streaks would develop wherein cooling fluid film protection would fail. This, in turn, could result in local damage to or destruction of the combustion chamber liner;
  • FIGSJZ and 3 In order to strengthen the overhanging lip 72' against such stresses, as well as relieve the residual stresses therein, there have been provided in the past a plurality of reinforcing dimples within the material of the lip itself.
  • the dimples are designated 80, and can be seen to be substantially conical in the cross section, originating in a point 82 at their upstream ends andterminating in a substantially circular cross section 84 at theirdownstream ends.
  • the dimples are formed as depressions in the .lip material itself, with the depression extending in the direction from the lip surface toward the associated downstream liner segment 62 (and, hence, away from the centerline of the combustion zone).
  • the lateral extremities86 and 88 of an individual dimple comprise the intersections between the conical cross section of the depression and the plane of lip 72. These lateral extremities divergein the downsteam direction with respect to any individual dimple;
  • the dimple configurations create an aerodynamic disturbance which reduces film cooling effectiveness downstream of the dimple.
  • each dimple designated 90, includes a substantially radially depressed portion 92, this portion having a substantially larger circumfer-
  • the upstream extremity 98 of each dimple is substantially linear in the circumferential direction with respect to the lip, while the 'downstreamfextremity is substantially arcuate describing an arc 100 in the cross pressed portion 92 from. greater width upstream at 94 to smaller downstream width-96.
  • the lateral extremities 102 and 104 -of individual dimples converge in the downstream direction while the adjacent extremities of adjacent dimples diverge in the downstream direction.
  • the reinforcing dimple of the present invention serves to increase the. life of the overhanging lip and, hence, the reliability of the associated combustor liner over extended periods of applicationQOne facet of this improvement rests in a substantial increase in overall fatigue strength due to reduction in the localized stresses in the lip structure proximate the dimple.
  • This improvement is dramatically exemplified in FIG. 6 which displaysa graphical representation of stress versus position over various stations through dimples according to the prior art and present invention. It can be appreciated frornthis Figure that localstressesare generally maintained at levels approximately one-half of the magnitude of the prior art dimple by means of utilization of the present invention. Experiments have illustrated that this stress reductionresults in a life gain of on a ratio of 15 to 1 over the prior art cooling slot.
  • An additional benefit of the present invention relates to the efiiciency with which the cooling fluid operates to remove heat from the combustion chamber liner.
  • This can be expressed in terms of cooling film effectiveness, 1 which is depicted in graphical form in FIG. 7 and plotted versus axial distance downstream of the dimple along'the liner surface.
  • the figure illustrates a comparison in efficiency between the prior art dimple andthe dimple according to the present invention, both directly downstream of individual dimples and inthe areas between dimples downstream thereof.
  • a marked improvement infilm barrier effectiveness is achieved by utilization of the dimples according to the present invention. This improvement can be traced'to aerodynamic improvements of the present dimple configuration which cause the film flow streamlines to converge and, thus, minimize disturbance or wake effects downstream of the dimples.
  • a cooling fluid plenum In a combustor liner film cooling slot of the variety including an annular space extending circumferentially of the liner and defined between overlapping portions of telescoping liner segments, a cooling fluid plenum.
  • said depresssed portion further includes lateral xtremities, and said lateral extremities substantially converge in the downstream direction.
  • each of said dimples departs from the surface of said lip to a first depth proximate its uppressed portion extends radially toward said liner.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Spray-Type Burners (AREA)
US00346595A 1973-03-30 1973-03-30 Combustion liner cooling slot stabilizing dimple Expired - Lifetime US3826082A (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
US00346595A US3826082A (en) 1973-03-30 1973-03-30 Combustion liner cooling slot stabilizing dimple
CA192,405A CA996761A (en) 1973-03-30 1974-02-13 Combustion liner cooling slot stabilizing dimple
DE2414376A DE2414376A1 (de) 1973-03-30 1974-03-26 Kuehlschlitzanordnung fuer einen brennkammereinsatz
IT49727/74A IT1003927B (it) 1973-03-30 1974-03-27 Ondulazione di stabilizzazione per cava di raffreddamento di una cami cia di camera di combustione parti colarmente per turbomotori a gas
BE142678A BE813091A (fr) 1973-03-30 1974-03-29 Fente de refroidissement de chemise de chambre de combustion
JP49034778A JPS5025918A (enrdf_load_stackoverflow) 1973-03-30 1974-03-29
FR7411194A FR2223554A1 (enrdf_load_stackoverflow) 1973-03-30 1974-03-29

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US00346595A US3826082A (en) 1973-03-30 1973-03-30 Combustion liner cooling slot stabilizing dimple

Publications (1)

Publication Number Publication Date
US3826082A true US3826082A (en) 1974-07-30

Family

ID=23360142

Family Applications (1)

Application Number Title Priority Date Filing Date
US00346595A Expired - Lifetime US3826082A (en) 1973-03-30 1973-03-30 Combustion liner cooling slot stabilizing dimple

Country Status (7)

Country Link
US (1) US3826082A (enrdf_load_stackoverflow)
JP (1) JPS5025918A (enrdf_load_stackoverflow)
BE (1) BE813091A (enrdf_load_stackoverflow)
CA (1) CA996761A (enrdf_load_stackoverflow)
DE (1) DE2414376A1 (enrdf_load_stackoverflow)
FR (1) FR2223554A1 (enrdf_load_stackoverflow)
IT (1) IT1003927B (enrdf_load_stackoverflow)

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2657471A1 (de) * 1975-12-22 1977-06-30 Gen Electric Kuehlschlitzanordnung im mantel eines brenners
US4109459A (en) * 1974-07-19 1978-08-29 General Electric Company Double walled impingement cooled combustor
DE2949473A1 (de) * 1978-12-11 1980-06-19 Gen Electric Brennerauskleidungsschlitz mit gekuehlten streben
US4329848A (en) * 1979-03-01 1982-05-18 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Cooling of combustion chamber walls using a film of air
US4688310A (en) * 1983-12-19 1987-08-25 General Electric Company Fabricated liner article and method
US4723413A (en) * 1985-11-19 1988-02-09 MTU Munuch, GmbH Reverse flow combustion chamber, especially reverse flow ring combustion chamber, for gas turbine propulsion units, with at least one flame tube wall film-cooling arrangement
US4833881A (en) * 1984-12-17 1989-05-30 General Electric Company Gas turbine engine augmentor
EP0434361A1 (en) * 1989-12-22 1991-06-26 Hitachi, Ltd. Combustion apparatus and combustion method therein
DE4142413A1 (de) * 1991-11-08 1993-05-19 Bmw Rolls Royce Gmbh Brennkammergehaeuse einer gasturbine
US5259182A (en) * 1989-12-22 1993-11-09 Hitachi, Ltd. Combustion apparatus and combustion method therein
US5331803A (en) * 1989-07-24 1994-07-26 Sundstrand Corporation Method of obtaining a desired temperature profile in a turbine engine and turbine engine incorporating the same
RU2319075C1 (ru) * 2006-05-10 2008-03-10 Открытое акционерное общество "Климов" Жаровая труба камеры сгорания
US20100107645A1 (en) * 2008-10-31 2010-05-06 General Electric Company Combustor liner cooling flow disseminator and related method
US20150198335A1 (en) * 2014-01-16 2015-07-16 Doosan Heavy Industries & Construction Co., Ltd. Liner, flow sleeve and gas turbine combustor each having cooling sleeve
US20150323183A1 (en) * 2014-05-08 2015-11-12 United Technologies Corporation Case with integral heat shielding
US20180156459A1 (en) * 2016-02-01 2018-06-07 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with wall contouring
US20190293292A1 (en) * 2016-05-23 2019-09-26 Mitsubishi Hitachi Power Systems, Ltd. Combustor and gas turbine
RU205407U1 (ru) * 2020-12-08 2021-07-13 Акционерное общество "Объединенная двигателестроительная корпорация" (АО "ОДК") Жаровая труба с компенсационными щелями
US20230003383A1 (en) * 2020-03-23 2023-01-05 Mitsubishi Heavy Industries, Ltd. Combustor and gas turbine provided with same
US11994291B2 (en) 2022-07-21 2024-05-28 General Electric Company Performance factor for a combustion liner

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CH620372A5 (en) * 1976-02-05 1980-11-28 Metaux Speciaux Sa Process for the treatment of spent catalysts with a view to the recovery of molybdenum
JPS6115418Y2 (enrdf_load_stackoverflow) * 1977-09-12 1986-05-13
FR2510541A2 (fr) * 1977-09-30 1983-02-04 Pechiney Ugine Kuhlmann Uran Procede de valorisation du molybdene a partir de solutions molybdeniferes contenant des carbonate, sulfate, hydroxyde ou hydrogenocarbonate alcalins, ainsi que, eventuellement, de l'uranium
US4380906A (en) * 1981-01-22 1983-04-26 United Technologies Corporation Combustion liner cooling scheme
US4529358A (en) * 1984-02-15 1985-07-16 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Vortex generating flow passage design for increased film cooling effectiveness
JPH0637976B2 (ja) * 1985-04-05 1994-05-18 工業技術院長 ガスタービン用燃焼器の尾筒の製造方法

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3307354A (en) * 1965-10-01 1967-03-07 Gen Electric Cooling structure for overlapped panels
US3589127A (en) * 1969-02-04 1971-06-29 Gen Electric Combustion apparatus
US3745766A (en) * 1971-10-26 1973-07-17 Avco Corp Variable geometry for controlling the flow of air to a combustor
US3751910A (en) * 1972-02-25 1973-08-14 Gen Motors Corp Combustion liner

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3307354A (en) * 1965-10-01 1967-03-07 Gen Electric Cooling structure for overlapped panels
US3589127A (en) * 1969-02-04 1971-06-29 Gen Electric Combustion apparatus
US3745766A (en) * 1971-10-26 1973-07-17 Avco Corp Variable geometry for controlling the flow of air to a combustor
US3751910A (en) * 1972-02-25 1973-08-14 Gen Motors Corp Combustion liner

Cited By (31)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4109459A (en) * 1974-07-19 1978-08-29 General Electric Company Double walled impingement cooled combustor
DE2657471A1 (de) * 1975-12-22 1977-06-30 Gen Electric Kuehlschlitzanordnung im mantel eines brenners
FR2336634A1 (fr) * 1975-12-22 1977-07-22 Gen Electric Fente de refroidissement par film pour chemise de chambre de combustion
US4050241A (en) * 1975-12-22 1977-09-27 General Electric Company Stabilizing dimple for combustion liner cooling slot
DE2949473A1 (de) * 1978-12-11 1980-06-19 Gen Electric Brennerauskleidungsschlitz mit gekuehlten streben
FR2444231A1 (fr) * 1978-12-11 1980-07-11 Gen Electric Chemise de chambre de combustion
US4259842A (en) * 1978-12-11 1981-04-07 General Electric Company Combustor liner slot with cooled props
US4329848A (en) * 1979-03-01 1982-05-18 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Cooling of combustion chamber walls using a film of air
US4688310A (en) * 1983-12-19 1987-08-25 General Electric Company Fabricated liner article and method
US4833881A (en) * 1984-12-17 1989-05-30 General Electric Company Gas turbine engine augmentor
US4723413A (en) * 1985-11-19 1988-02-09 MTU Munuch, GmbH Reverse flow combustion chamber, especially reverse flow ring combustion chamber, for gas turbine propulsion units, with at least one flame tube wall film-cooling arrangement
US5331803A (en) * 1989-07-24 1994-07-26 Sundstrand Corporation Method of obtaining a desired temperature profile in a turbine engine and turbine engine incorporating the same
US5259182A (en) * 1989-12-22 1993-11-09 Hitachi, Ltd. Combustion apparatus and combustion method therein
EP0434361A1 (en) * 1989-12-22 1991-06-26 Hitachi, Ltd. Combustion apparatus and combustion method therein
DE4142413A1 (de) * 1991-11-08 1993-05-19 Bmw Rolls Royce Gmbh Brennkammergehaeuse einer gasturbine
RU2319075C1 (ru) * 2006-05-10 2008-03-10 Открытое акционерное общество "Климов" Жаровая труба камеры сгорания
RU2319075C9 (ru) * 2006-05-10 2008-05-27 Открытое акционерное общество "Климов" Жаровая труба камеры сгорания
US20100107645A1 (en) * 2008-10-31 2010-05-06 General Electric Company Combustor liner cooling flow disseminator and related method
CN101725977A (zh) * 2008-10-31 2010-06-09 通用电气公司 燃烧器衬套冷却流散布件及相关方法
EP2182286A3 (en) * 2008-10-31 2014-04-30 General Electric Company Combustor Liner Cooling Flow Disseminator and Related Method
US20150198335A1 (en) * 2014-01-16 2015-07-16 Doosan Heavy Industries & Construction Co., Ltd. Liner, flow sleeve and gas turbine combustor each having cooling sleeve
US10094573B2 (en) * 2014-01-16 2018-10-09 DOOSAN Heavy Industries Construction Co., LTD Liner, flow sleeve and gas turbine combustor each having cooling sleeve
US20150323183A1 (en) * 2014-05-08 2015-11-12 United Technologies Corporation Case with integral heat shielding
US10012389B2 (en) * 2014-05-08 2018-07-03 United Technologies Corporation Case with integral heat shielding
US20180156459A1 (en) * 2016-02-01 2018-06-07 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with wall contouring
US10670270B2 (en) * 2016-02-01 2020-06-02 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with wall contouring
US20190293292A1 (en) * 2016-05-23 2019-09-26 Mitsubishi Hitachi Power Systems, Ltd. Combustor and gas turbine
US11085642B2 (en) * 2016-05-23 2021-08-10 Mitsubishi Power, Ltd. Combustor with radially varying leading end portion of basket and gas turbine
US20230003383A1 (en) * 2020-03-23 2023-01-05 Mitsubishi Heavy Industries, Ltd. Combustor and gas turbine provided with same
RU205407U1 (ru) * 2020-12-08 2021-07-13 Акционерное общество "Объединенная двигателестроительная корпорация" (АО "ОДК") Жаровая труба с компенсационными щелями
US11994291B2 (en) 2022-07-21 2024-05-28 General Electric Company Performance factor for a combustion liner

Also Published As

Publication number Publication date
CA996761A (en) 1976-09-14
BE813091A (fr) 1974-07-15
FR2223554A1 (enrdf_load_stackoverflow) 1974-10-25
IT1003927B (it) 1976-06-10
JPS5025918A (enrdf_load_stackoverflow) 1975-03-18
DE2414376A1 (de) 1974-10-17

Similar Documents

Publication Publication Date Title
US3826082A (en) Combustion liner cooling slot stabilizing dimple
US4622821A (en) Combustion liner for a gas turbine engine
US4555901A (en) Combustion chamber construction
US4380906A (en) Combustion liner cooling scheme
US5282721A (en) Passive clearance system for turbine blades
US4023350A (en) Exhaust case for a turbine machine
US4353679A (en) Fluid-cooled element
US3365172A (en) Air cooled shroud seal
US4050241A (en) Stabilizing dimple for combustion liner cooling slot
US3220697A (en) Hollow turbine or compressor vane
US3844343A (en) Impingement-convective cooling system
US3793827A (en) Stiffener for combustor liner
US3800864A (en) Pin-fin cooling system
US4920742A (en) Heat shield for gas turbine engine frame
US4013376A (en) Coolable blade tip shroud
US7147432B2 (en) Turbine shroud asymmetrical cooling elements
US4733538A (en) Combustion selective temperature dilution
US4739621A (en) Cooling scheme for combustor vane interface
EP1726811B1 (en) System and method for cooling lateral edge regions of a divergent seal of an axisymmetric nozzle
EP0086667B1 (en) Improved low smoke combustor for land based combustion turbines
US5542246A (en) Bulkhead cooling fairing
US3978664A (en) Gas turbine engine diffuser
US4392355A (en) Combustion liner
US4955192A (en) Containment ring for radial inflow turbine
US5080284A (en) Cooling system for the trailing edge of a liner