US3706203A - Wall structure for a gas turbine engine - Google Patents

Wall structure for a gas turbine engine Download PDF

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Publication number
US3706203A
US3706203A US85629A US3706203DA US3706203A US 3706203 A US3706203 A US 3706203A US 85629 A US85629 A US 85629A US 3706203D A US3706203D A US 3706203DA US 3706203 A US3706203 A US 3706203A
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United States
Prior art keywords
liner
wall
combustion
chamber
passages
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Expired - Lifetime
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US85629A
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English (en)
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Perry Goldberg
Irwin Segalman
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RTX Corp
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United Aircraft Corp
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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
    • F23R3/02—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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a heat resistant wall construction, and more particularly to a wall construction which has particular utility in the hot temperature environment of a gas turbine engine.
  • the combustion chambers utilize a cooled combustion chamber liner or wall construction to limit the temperature of the main load carrying structures of the engine such as the inner and outer walls of the burner section of the engine.
  • this liner also serves to control the fuel-air distribution in the primary combustion zone and to permit controlled mixing of unburned air with the combustion products to achieve desired turbine inlet temperature profiles.
  • the cooling for the liner or wall construction which actually defines a combustion zone within which the combustion or burning occurs, is intended to limit the temperature of the liner wall itself and thereby the temperature of the surrounding main load carrying structural members.
  • Conventional combustion chamber liners have a 'number of drawbacks. For example, these constructions require large quantities of cooling air hence reducing the air available for combustion and/or dilution in the main burners. This limits the combustor performance in terms of temperature rise capability and quality of exit temperature profile. It can also result in longer combustion chambers or increased combustor pressure loss to achieve the desired turbine inlet temperature profiles with the reduced dilution air. These obviously result in weight and performance penalties for the gas turbine engine.
  • the present invention provides a burner construction which avoids the principal problems encountered by the prior art in a construction in which the cooling air requirements are reduced by approximately 50 percent. Additionally, the construction provides a liner or a wall structure which is substantially free from severe thermal gradients, therefore, improving the life characteristics of the burner.
  • the burner or combustion chamber includes a liner or wall construction which comprises a pair of radially spaced walls.
  • the radial space is actually a flow passageway which extends over the entire length of a liner wall segment.
  • the flow passageway in turn comprises a plurality of relatively small diameter flow channels, the flow channels being formed by either openings or through the use of a corrugated strip member.
  • a plurality of double wall segments are joined to one another to form sets of flow channels extending over the entire length of each of the wall segments, with the length of each segment controlled with respect to the temperature in the adjacent portion of the combustion area.
  • Each of the flow channels has a relatively large length-to-diameter ratio.
  • the inner wall of the liner actually confines or is the boundary, member for the hot combustion gases, and it is this wall in the liner construction which has to be maintained at an acceptance operating temperature.
  • a cooling stream with a temperature rela- I tively cooler than the combustion gases, is introduced into the flow channels.
  • the frictional pressure losses of the cooling stream in each of the flow channels can be established and controlled, and thus the operating temperature of the inner wall can be maintained at an acceptable level. Additionally this is achieved with an amount of cooling flow which is approximately 50 percent less than other bumer construction.
  • Another feature of the present invention is that since the flow channels extend over the entire length, and since the control of the cooling flow can be accurately controlled, the temperature of the inner wall is maintained at a uniform level. Therefore the adverse thermal gradient and stress loads are substantially eliminated.
  • FIG. 1 is a partial cross-sectional view of the burner apparatus, the burner apparatus including a liner wall means utilizing the construction of the present invention.
  • FIG. 2 is a sectional view taken substantially along line 2-2.
  • FIG. 3 is a sectional view similar to FIG. 2 showing a modification.
  • FIG. 4 is a plot showing diagrammatically the relation of burner temperature, liner temperature, and length of wall segments.
  • the invention is shown in a burner can or liner for a combustion chamber which is located between the compressor and turbine of a gas turbine power plant and in which fuel is burned in the high pressure gas discharged from the compressor to provide a hot gas under pressure for expansion through the turbine.
  • a power plant to which this type of combustion chamber is applicable is disclosed for example in the Savin U.S. Pat. No. 2,747,367.
  • the combustion chamber is of a can-annular type only one can being shown. It is understood that any type of combustion chamber may be employed whether it be a can-annular type or an annular type.
  • the can-annular combustion chamber has an inner cylindrical wall 16 and an outer wall 17 both attached at their upstream ends to a diffuser 18. Within the diffuser is mounted a fuel nozzle 19 for each burner can 20 or liner.
  • the liner or can includes a domeshaped head 22 at the inlet end with an opening therein to receive the fuel nozzle. Combustion occurs within this burner can or liner and the products thereof discharge from the open downstream end 24 into the turbine.
  • High pressure air from the compressor is discharged into and through the diffuser 18 and flows into the combustion chamber. A portion of this air enters the liner or can20 through a swirler 26 around the fuel nozzle and also through a plurality of combustion holes 28 which are positioned within the liner wall.
  • This liner is made up of a plurality of rings or segments 20a, 20b, 20c, 20d, and 20e which are attached to one another and through the walls of which cooling air flows for the purpose of maintaining the wall temperature within the desired limits. It is essential that the liner and the surrounding structural members of the engine be maintained within acceptable operating temperature ranges and to this extent the liner 20 has a cooled wall of novel construction.
  • this wall is double thickness and has an inner wall element 40 and radially spaced therefrom, an outer wall element 42.
  • the inner wall element 40 is directly exposed to the combustion gases within the liner or can and because of the temperature of these gases it becomes necessary to cool the wall to a suitable operating temperature to prevent damage to this wall during operation of the engine.
  • Cooling flow passages 44 are provided between the wall 40 and 42 by circumferentially spaced, longitudinally extending ribs 46 integral with and projecting outwardly from the inner wall and into contact with the outer wall, the spaces between the ribs defining the flow passages.
  • the burner can segments 20a, 20 b, 20c, 20d, and 20 e are each of the double thickness described and with the downstream end 52 of each segment attached to and positioned within the upstream end 54 of the next adjacent liner segment.
  • the arrangement of these segments is such that the cooling air enters the upstream end of the passages 44 from outside of the liner, flows through the passages and is discharged into the space inside of the liner in a direction substantially parallel to the liner and directly within the liner.
  • This arrangement is shown in FIG. 1 in which each of the liner sections tapers slightly from the upstream end to the downstream end as shown and the purpose of this is so that the successive liner sections will telescope one with respect with another.
  • the inner wall element 40 of the first liner section extends over the edge of the dome 22 for. the burner can and the downstream end of the outer wall element 42 of this liner section fits within the upstream end of the inner wall element 40 of the next liner section. Also as shown in FIG. 1, the inner wall element 40 extends forwardly somewhat beyond the passages within the liner so that this wall is exposed to permit welding or other attachment to the underlying end of the outer wall of the adjacent liner segment.
  • adjacent liner sections may be suitably attached one to another by welding the overlapping and contacting inner wall of one segment with the outer wall of the adjacent segment.
  • the liner segments vary in length depending upon the temperature of the combustion gases within the burner can in order that the wall temperature of the liner segments may remain substantially constant and at no time reach a point above that at which the burner can may operate successfully without damage to the material of the liner wall.
  • the first liner segment is made only 1 inch in length.
  • the second segment is one and three eighths of an inch, the third segment is 1% and the successive segments are longer since at this time the gas temperature within the burner can is diminishing as shown and less cooling is neccessary to accomplish the desired result of maintaining a workable temperature for the liner wall.
  • the wall of the burner can consists of an inner wall element 60, an outer parallel wall element 62 radially spaced therefrom, and a corrugated sheet 64 positioned therebetween and bonded to both to hold them in spaced relation and to define a plurality of longitudinally extending circumferentially spaced passages 66.
  • the effective diameter of these passages for the particular burner can shown which has a diameter of approximately 6.5 inches was 0.04 inches.
  • These double thickness wall elements are made up in the same manner as shown in FIG. 1 with the outer wall extending beyond the passages at the downstream end and with the inner wall extending beyond the passages at the upstream end to provide overlapping flanges by which successive segments may be secured together.
  • FIG. 2 or FIG. 3 the additional coolant side surface area as a consequence of the dividers between individual coolant channels, significantly enhances the thermal effectiveness of the wall construction and thereby reduces the cooling air requirement.
  • each segment may have a small taper from end to end so that when the several seg ments are secured together, the resulting can will have a substantially constant diameter from end to end.
  • this plot represents a burner can that was built and tested and shows the temperature variation within the can and the temperatures at which the successive sections of the liner operated.
  • the particular wall construction of the burner can used was the arrangement of FIG. 3 above described, and was of such dimension that the effective diameter of each of the cooling passages was 0.04 inch, so that the length-todiameter ratio for the first section was 25.
  • the successive burner can segments lengthening as the temperature within the can decreases the amount of cooling air needed is materially reduced, thereby minimizing the amount of cooling air required.
  • the coolant side heat transfer is significantly enhanced.
  • the smaller the area of the channel the greater the surface area of the channel that is exposed to the air passing through the channel, and thereby, the greater 'cooling effect.
  • the ability-to cool the structure with a minimum of cooling air is enhanced by significantly increasing the heat transfer coefficients and a significant increase in coolant side surface area.
  • the liner or burner can segments extend circumferentially but obviously the flow passages are longitudinally or axially of the'combustion chamber, and the length of each segment determines the length of the flow passages in that segment.
  • the liner segments increase in length toward the downstream end of the liner or burner can and thus the length of segments increases as the operating temperature within the can decreases.
  • the essential feature is to provide only enough cooling of the wall to keep the temperature of the wall from exceeding the safe operating temperature, normally established for the particular alloy used as 1,600F in the particular arrangement shown. This has been accomplished in the construction shown as evidenced by the chart of FIG. 4.
  • a liner for a combustion chamber for use in a gas turbine engine in which combustion occurs in a gas stream moving axially through the combustion chamber said liner including a plurality of liner segments arranged circumferentially, each segment having spaced inner and outer wall elements and means interconnecting the elements and defining closely spaced longitudinally extending flow passages therebetween, the outer wall of one segment being connected to and overlying the inner wall of the next adjacent downstream segment to cause cooling air to flow into the upstream ends of the passages from the chamber space outside the liner and to enter the space inside the liner at the downstream ends of said passages, successive liner segments increasing in length in a downstream direction as the operating temperature within the liner decreases.
  • annular combustion chamber for a gas turbine engine in which combustion occurs in a stream of gas moving axially through said chamber, said chamber having inner and outer walls forming an annular space therebetween and a liner spaced from one of the said walls to define between said one wall and the liner a passage for cooling air and on the other side of the liner a combustion space, said liner comprising a plurality of segments extending in a circumferential direction within the chamber and located in side-by-side relation axially of the chamber and secured together, each segment having spaced inner and outer wall elements and spacer means between said elements serving to hold said elements in spaced relation and forming closely spaced parallel passages between said wall elements extending in an axial direction, the inner wall element of one liner se ment overlying and secured to the outer wa l elemen of the ad acen upstream element and said successive segments increasing in length toward the downstream end of the combustion chamber for increasing the lengths of the flow passages near the areas of decreasing temperature within the combustion space, each of said passages being

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Heat-Exchange Devices With Radiators And Conduit Assemblies (AREA)
US85629A 1970-10-30 1970-10-30 Wall structure for a gas turbine engine Expired - Lifetime US3706203A (en)

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US8562970A 1970-10-30 1970-10-30

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US (1) US3706203A (fr)
AU (1) AU3490971A (fr)
BE (1) BE774560A (fr)
CA (1) CA938797A (fr)
CH (1) CH534847A (fr)
DE (1) DE2147135A1 (fr)
FR (1) FR2111931B1 (fr)
GB (1) GB1314666A (fr)
IL (1) IL37773A (fr)
NL (1) NL7113326A (fr)

Cited By (45)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3938323A (en) * 1971-12-15 1976-02-17 Phillips Petroleum Company Gas turbine combustor with controlled fuel mixing
US3955361A (en) * 1971-12-15 1976-05-11 Phillips Petroleum Company Gas turbine combustor with controlled fuel mixing
US4012902A (en) * 1974-03-29 1977-03-22 Phillips Petroleum Company Method of operating a gas turbine combustor having an independent airstream to remove heat from the primary combustion zone
US4184326A (en) * 1975-12-05 1980-01-22 United Technologies Corporation Louver construction for liner of gas turbine engine combustor
FR2479951A1 (fr) * 1980-04-02 1981-10-09 United Technologies Corp Garniture interieure de chambre de combustion
FR2479900A1 (fr) * 1980-04-02 1981-10-09 United Technologies Corp Garniture interieure de chambre de combustion
FR2479901A1 (fr) * 1980-04-02 1981-10-09 United Technologies Corp Garniture interieure de chambre de combustion pour une turbine a gaz
US4333216A (en) * 1981-03-23 1982-06-08 United Technologies Corporation Method for manufacturing a sandwich panel structure
US4407205A (en) * 1982-04-30 1983-10-04 Beaufrere Albert H Regeneratively cooled coal combustor/gasifier with integral dry ash removal
US4619604A (en) * 1983-06-30 1986-10-28 Carrier Corporation Flame radiator structure
US5323601A (en) * 1992-12-21 1994-06-28 United Technologies Corporation Individually removable combustor liner panel for a gas turbine engine
US5327727A (en) * 1993-04-05 1994-07-12 General Electric Company Micro-grooved heat transfer combustor wall
US5329773A (en) * 1989-08-31 1994-07-19 Alliedsignal Inc. Turbine combustor cooling system
US6408628B1 (en) * 1999-11-06 2002-06-25 Rolls-Royce Plc Wall elements for gas turbine engine combustors
US6868675B1 (en) * 2004-01-09 2005-03-22 Honeywell International Inc. Apparatus and method for controlling combustor liner carbon formation
US20060010874A1 (en) * 2004-07-15 2006-01-19 Intile John C Cooling aft end of a combustion liner
US20060096293A1 (en) * 2004-11-08 2006-05-11 United Technologies Corporation Pulsed combustion engine
US20060277921A1 (en) * 2005-06-10 2006-12-14 Pratt & Whitney Canada Corp. Gas turbine engine combustor with improved cooling
RU2292516C2 (ru) * 2001-11-30 2007-01-27 Пауэ Системс Мфг. Ллс Блок камеры сгорания и способ охлаждения трубки вентури в этом блоке
US20090277180A1 (en) * 2008-05-07 2009-11-12 Kam-Kei Lam Combustor dynamic attenuation and cooling arrangement
US20100170259A1 (en) * 2009-01-07 2010-07-08 Huffman Marcus B Method and apparatus to enhance transition duct cooling in a gas turbine engine
US20110232299A1 (en) * 2010-03-25 2011-09-29 Sergey Aleksandrovich Stryapunin Impingement structures for cooling systems
US20120003595A1 (en) * 2009-09-29 2012-01-05 Honeywell International Inc. High turn down low nox burner
WO2012112514A1 (fr) * 2011-02-14 2012-08-23 Icr Turbine Engine Corporation Écran contre le rayonnement pour une chambre de combustion d'une turbine à gaz
US8307654B1 (en) * 2009-09-21 2012-11-13 Florida Turbine Technologies, Inc. Transition duct with spiral finned cooling passage
US8402764B1 (en) * 2009-09-21 2013-03-26 Florida Turbine Technologies, Inc. Transition duct with spiral cooling channels
US20130174558A1 (en) * 2011-08-11 2013-07-11 General Electric Company System for injecting fuel in a gas turbine engine
US8499874B2 (en) 2009-05-12 2013-08-06 Icr Turbine Engine Corporation Gas turbine energy storage and conversion system
CN103547866A (zh) * 2011-03-29 2014-01-29 西门子能量股份有限公司 涡轮燃烧系统衬垫
US8669670B2 (en) 2010-09-03 2014-03-11 Icr Turbine Engine Corporation Gas turbine engine configurations
EP1795806A3 (fr) * 2005-12-06 2014-05-28 Deutsches Zentrum für Luft- und Raumfahrt e.V. Chambre chaude
US8866334B2 (en) 2010-03-02 2014-10-21 Icr Turbine Engine Corporation Dispatchable power from a renewable energy facility
US8984895B2 (en) 2010-07-09 2015-03-24 Icr Turbine Engine Corporation Metallic ceramic spool for a gas turbine engine
US9051873B2 (en) 2011-05-20 2015-06-09 Icr Turbine Engine Corporation Ceramic-to-metal turbine shaft attachment
CN105318356A (zh) * 2014-07-21 2016-02-10 北京航天动力研究所 一种大深宽比变截面换热通道
US20170176005A1 (en) * 2015-12-17 2017-06-22 Rolls-Royce Plc Combustion chamber
US20180073390A1 (en) * 2016-09-13 2018-03-15 Rolls-Royce Corporation Additively deposited gas turbine engine cooling component
US10094288B2 (en) 2012-07-24 2018-10-09 Icr Turbine Engine Corporation Ceramic-to-metal turbine volute attachment for a gas turbine engine
US10450871B2 (en) 2015-02-26 2019-10-22 Rolls-Royce Corporation Repair of dual walled metallic components using directed energy deposition material addition
US10478920B2 (en) 2014-09-29 2019-11-19 Rolls-Royce Corporation Dual wall components for gas turbine engines
US10766105B2 (en) 2015-02-26 2020-09-08 Rolls-Royce Corporation Repair of dual walled metallic components using braze material
RU201848U1 (ru) * 2020-08-12 2021-01-15 федеральное государственное бюджетное образовательное учреждение высшего образования "Ульяновский государственный технический университет" Камера сгорания газотурбинного двигателя с активной зоной охлаждения
US11480337B2 (en) 2019-11-26 2022-10-25 Collins Engine Nozzles, Inc. Fuel injection for integral combustor and turbine vane
US12036627B2 (en) 2018-03-08 2024-07-16 Rolls-Royce Corporation Techniques and assemblies for joining components
EP4596971A4 (fr) * 2022-11-07 2025-12-31 Mitsubishi Heavy Ind Aero Engines Ltd Corps cylindrique

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH0648095B2 (ja) * 1985-04-18 1994-06-22 石川島播磨重工業株式会社 ガスタ−ビン燃焼器等のライナ−冷却構造
US4642993A (en) * 1985-04-29 1987-02-17 Avco Corporation Combustor liner wall

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2458066A (en) * 1944-07-20 1949-01-04 American Locomotive Co Combustion chamber
US2617255A (en) * 1947-05-12 1952-11-11 Bbc Brown Boveri & Cie Combustion chamber for a gas turbine
US2775094A (en) * 1953-12-03 1956-12-25 Gen Electric End cap for fluid fuel combustor
US3154914A (en) * 1959-12-12 1964-11-03 Bolkow Entwicklungen Kg Rocket engine construction
US3545202A (en) * 1969-04-02 1970-12-08 United Aircraft Corp Wall structure and combustion holes for a gas turbine engine
US3572031A (en) * 1969-07-11 1971-03-23 United Aircraft Corp Variable area cooling passages for gas turbine burners

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2458066A (en) * 1944-07-20 1949-01-04 American Locomotive Co Combustion chamber
US2617255A (en) * 1947-05-12 1952-11-11 Bbc Brown Boveri & Cie Combustion chamber for a gas turbine
US2775094A (en) * 1953-12-03 1956-12-25 Gen Electric End cap for fluid fuel combustor
US3154914A (en) * 1959-12-12 1964-11-03 Bolkow Entwicklungen Kg Rocket engine construction
US3545202A (en) * 1969-04-02 1970-12-08 United Aircraft Corp Wall structure and combustion holes for a gas turbine engine
US3572031A (en) * 1969-07-11 1971-03-23 United Aircraft Corp Variable area cooling passages for gas turbine burners

Cited By (59)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3938323A (en) * 1971-12-15 1976-02-17 Phillips Petroleum Company Gas turbine combustor with controlled fuel mixing
US3955361A (en) * 1971-12-15 1976-05-11 Phillips Petroleum Company Gas turbine combustor with controlled fuel mixing
US4012902A (en) * 1974-03-29 1977-03-22 Phillips Petroleum Company Method of operating a gas turbine combustor having an independent airstream to remove heat from the primary combustion zone
US4184326A (en) * 1975-12-05 1980-01-22 United Technologies Corporation Louver construction for liner of gas turbine engine combustor
DE3113379A1 (de) * 1980-04-02 1982-05-06 United Technologies Corp., 06101 Hartford, Conn. Brennereinsatz fuer ein gasturbinentriebwerk
FR2479900A1 (fr) * 1980-04-02 1981-10-09 United Technologies Corp Garniture interieure de chambre de combustion
FR2479901A1 (fr) * 1980-04-02 1981-10-09 United Technologies Corp Garniture interieure de chambre de combustion pour une turbine a gaz
DE3113380A1 (de) * 1980-04-02 1982-04-08 United Technologies Corp., 06101 Hartford, Conn. Einsatz fuer den brenner eines gasturbinentriebwerks
FR2479951A1 (fr) * 1980-04-02 1981-10-09 United Technologies Corp Garniture interieure de chambre de combustion
US4333216A (en) * 1981-03-23 1982-06-08 United Technologies Corporation Method for manufacturing a sandwich panel structure
US4407205A (en) * 1982-04-30 1983-10-04 Beaufrere Albert H Regeneratively cooled coal combustor/gasifier with integral dry ash removal
US4619604A (en) * 1983-06-30 1986-10-28 Carrier Corporation Flame radiator structure
US5329773A (en) * 1989-08-31 1994-07-19 Alliedsignal Inc. Turbine combustor cooling system
US5323601A (en) * 1992-12-21 1994-06-28 United Technologies Corporation Individually removable combustor liner panel for a gas turbine engine
EP0604021A1 (fr) * 1992-12-21 1994-06-29 United Technologies Corporation Paroi de protection individuellement enlevable d'une chambre de combustion pour une turbine à gaz
US5327727A (en) * 1993-04-05 1994-07-12 General Electric Company Micro-grooved heat transfer combustor wall
US6408628B1 (en) * 1999-11-06 2002-06-25 Rolls-Royce Plc Wall elements for gas turbine engine combustors
RU2292516C2 (ru) * 2001-11-30 2007-01-27 Пауэ Системс Мфг. Ллс Блок камеры сгорания и способ охлаждения трубки вентури в этом блоке
US6868675B1 (en) * 2004-01-09 2005-03-22 Honeywell International Inc. Apparatus and method for controlling combustor liner carbon formation
US20060010874A1 (en) * 2004-07-15 2006-01-19 Intile John C Cooling aft end of a combustion liner
US20060096293A1 (en) * 2004-11-08 2006-05-11 United Technologies Corporation Pulsed combustion engine
US7278256B2 (en) * 2004-11-08 2007-10-09 United Technologies Corporation Pulsed combustion engine
US20060277921A1 (en) * 2005-06-10 2006-12-14 Pratt & Whitney Canada Corp. Gas turbine engine combustor with improved cooling
US7509809B2 (en) * 2005-06-10 2009-03-31 Pratt & Whitney Canada Corp. Gas turbine engine combustor with improved cooling
EP1795806A3 (fr) * 2005-12-06 2014-05-28 Deutsches Zentrum für Luft- und Raumfahrt e.V. Chambre chaude
US20090277180A1 (en) * 2008-05-07 2009-11-12 Kam-Kei Lam Combustor dynamic attenuation and cooling arrangement
US9121610B2 (en) * 2008-05-07 2015-09-01 Siemens Aktiengesellschaft Combustor dynamic attenuation and cooling arrangement
US20100170259A1 (en) * 2009-01-07 2010-07-08 Huffman Marcus B Method and apparatus to enhance transition duct cooling in a gas turbine engine
US8549861B2 (en) * 2009-01-07 2013-10-08 General Electric Company Method and apparatus to enhance transition duct cooling in a gas turbine engine
US8708083B2 (en) 2009-05-12 2014-04-29 Icr Turbine Engine Corporation Gas turbine energy storage and conversion system
US8499874B2 (en) 2009-05-12 2013-08-06 Icr Turbine Engine Corporation Gas turbine energy storage and conversion system
US8307654B1 (en) * 2009-09-21 2012-11-13 Florida Turbine Technologies, Inc. Transition duct with spiral finned cooling passage
US8402764B1 (en) * 2009-09-21 2013-03-26 Florida Turbine Technologies, Inc. Transition duct with spiral cooling channels
US20120003595A1 (en) * 2009-09-29 2012-01-05 Honeywell International Inc. High turn down low nox burner
US8866334B2 (en) 2010-03-02 2014-10-21 Icr Turbine Engine Corporation Dispatchable power from a renewable energy facility
US20110232299A1 (en) * 2010-03-25 2011-09-29 Sergey Aleksandrovich Stryapunin Impingement structures for cooling systems
US8984895B2 (en) 2010-07-09 2015-03-24 Icr Turbine Engine Corporation Metallic ceramic spool for a gas turbine engine
US8669670B2 (en) 2010-09-03 2014-03-11 Icr Turbine Engine Corporation Gas turbine engine configurations
WO2012112514A1 (fr) * 2011-02-14 2012-08-23 Icr Turbine Engine Corporation Écran contre le rayonnement pour une chambre de combustion d'une turbine à gaz
CN103547866A (zh) * 2011-03-29 2014-01-29 西门子能量股份有限公司 涡轮燃烧系统衬垫
US9051873B2 (en) 2011-05-20 2015-06-09 Icr Turbine Engine Corporation Ceramic-to-metal turbine shaft attachment
US20130174558A1 (en) * 2011-08-11 2013-07-11 General Electric Company System for injecting fuel in a gas turbine engine
US9228499B2 (en) * 2011-08-11 2016-01-05 General Electric Company System for secondary fuel injection in a gas turbine engine
US10094288B2 (en) 2012-07-24 2018-10-09 Icr Turbine Engine Corporation Ceramic-to-metal turbine volute attachment for a gas turbine engine
CN105318356A (zh) * 2014-07-21 2016-02-10 北京航天动力研究所 一种大深宽比变截面换热通道
US10478920B2 (en) 2014-09-29 2019-11-19 Rolls-Royce Corporation Dual wall components for gas turbine engines
US10766105B2 (en) 2015-02-26 2020-09-08 Rolls-Royce Corporation Repair of dual walled metallic components using braze material
US10450871B2 (en) 2015-02-26 2019-10-22 Rolls-Royce Corporation Repair of dual walled metallic components using directed energy deposition material addition
US11731218B2 (en) 2015-02-26 2023-08-22 Rolls-Royce Corporation Repair of dual walled metallic components using braze material
US12157192B2 (en) 2015-02-26 2024-12-03 Rolls-Royce Corporation Repair of dual walled metallic components using braze material
US10533746B2 (en) * 2015-12-17 2020-01-14 Rolls-Royce Plc Combustion chamber with fences for directing cooling flow
US20170176005A1 (en) * 2015-12-17 2017-06-22 Rolls-Royce Plc Combustion chamber
US20180073390A1 (en) * 2016-09-13 2018-03-15 Rolls-Royce Corporation Additively deposited gas turbine engine cooling component
US11248491B2 (en) 2016-09-13 2022-02-15 Rolls-Royce Corporation Additively deposited gas turbine engine cooling component
US12036627B2 (en) 2018-03-08 2024-07-16 Rolls-Royce Corporation Techniques and assemblies for joining components
US11480337B2 (en) 2019-11-26 2022-10-25 Collins Engine Nozzles, Inc. Fuel injection for integral combustor and turbine vane
US11788723B2 (en) 2019-11-26 2023-10-17 Collins Engine Nozzles, Inc. Fuel injection for integral combustor and turbine vane
RU201848U1 (ru) * 2020-08-12 2021-01-15 федеральное государственное бюджетное образовательное учреждение высшего образования "Ульяновский государственный технический университет" Камера сгорания газотурбинного двигателя с активной зоной охлаждения
EP4596971A4 (fr) * 2022-11-07 2025-12-31 Mitsubishi Heavy Ind Aero Engines Ltd Corps cylindrique

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CA938797A (en) 1973-12-25
NL7113326A (fr) 1972-05-03
CH534847A (de) 1973-03-15
BE774560A (fr) 1972-02-14
AU3490971A (en) 1973-05-03
FR2111931A1 (fr) 1972-06-09
IL37773A (en) 1974-01-14
FR2111931B1 (fr) 1976-02-13
GB1314666A (en) 1973-04-26
DE2147135A1 (de) 1972-05-04
IL37773A0 (en) 1971-12-29

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