US3064425A - Combustion liner - Google Patents

Combustion liner Download PDF

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Publication number
US3064425A
US3064425A US844279A US84427959A US3064425A US 3064425 A US3064425 A US 3064425A US 844279 A US844279 A US 844279A US 84427959 A US84427959 A US 84427959A US 3064425 A US3064425 A US 3064425A
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United States
Prior art keywords
combustion
liner
air
section
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
US844279A
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English (en)
Inventor
Charles F Hayes
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.)
Motors Liquidation Co
Original Assignee
General Motors Corp
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 Motors Corp filed Critical General Motors Corp
Priority to US844279A priority Critical patent/US3064425A/en
Priority to GB31939/60A priority patent/GB891836A/en
Priority to FR840205A priority patent/FR1268931A/fr
Priority to CH1120260A priority patent/CH380444A/de
Application granted granted Critical
Publication of US3064425A publication Critical patent/US3064425A/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
    • F23R3/02—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

  • This invention relates to a combustion chamber construction, and more particularly, to a means for cooling the liner of a flame tube in a combustion section.
  • annular combustion cans In gas turbine engines having a cannular-type combustion section, a number of annular combustion cans, liners, or flame tubes, as they are commonly called, are generally spaced around the circumference of the combustion section.
  • One method of manufacturing these combustion cans is to telescope a number of axially aligned annular sections and spot weld them together at circumferentially spaced points along their overlapping edges.
  • hot spots often develop in the liner to the rear of the weld points or dimples due to an inadequate flow of cooling air to these areas thus causing an ultimate burn-out of the liner in this region.
  • This invention eliminates these faults by providing circumferential corrugations on the forward or upstream edge portion of each section cooperating with the rear overlapped edge portion of an adjacent section to provide fluid flow passages therebetween having shapes to direct the flow to the liner portions directly rearwardly or downstream of the weld points.
  • IGURE 1 is a side elevational view with parts broken away and in section of a portion of a gas turbine engine embodying the invention
  • FIGURE 2 is an enlarged portion of a detail of FIG- URE 1 with parts broken away and in section,
  • FIGURE 3 is an enlarged cross-sectional view of a detail taken on a plane indicated by and viewed in the direction of the arrows 33 of FIGURE 2,
  • FIGURES 4 and 5 are enlarged cross-sectional views of the FIG. 3 construction taken on planes indicated by and viewed in the direction of the arrows 4-4 and 55 of FIGURE 3,
  • FIGURE 6 is an enlarged perspective view of the details shown in FIGS. 35.
  • FIGURE 1 there is illustrated schematically therein a portion of a gas turbine engine 10 having a compressor 12 of the axial fiow type (only the later stages of which are shown), a diffuser section 14, an annular combustion section 16, and a turbine section 1% (partially shown).
  • Each of the cans has at its forward end a dome 28 adapted to cooperate with a fuel nozzle 30 of a conventional type secured to the engine casing in the diffuser section 14, and at its downwstream end a transition section 32 adapted to cooperate with the inlet to the turbine section 18. Furthermore, the cans are each pro vided with primary air inlet holes 36 and secondary and air dilution holes 38, as well as arcuate reinforcing members 40.
  • the cans 26 are also each provided with an opening 42 for the insertion therein of an igniter plug 44 secured at its opposite end to the engine casing. While the number of igniter plugs illustrated correspond to the number of combustion cans, generally one or two plugs are live, while the remaining ones are dummy plugs, with the propagation of the flame between the combustion cans being accomplished through the use of conventional crossover tubes.
  • the air discharged from the compressor 12 is directed into the difiuser 14 wherein the air velocity is reduced, the swirl component thereof is eliminated, and the dynamic pressure energy is changed to static pressure energy to present the air to the dome of the combustion can in a uniformly distributed fashion so that it will pass thereinto, be mixed with the fuel spray supplied through the fuel nozzle 30, ignited and burned, and pass into the turbine inlet to drive the turbine (not shown), which in turn drives the turbine shaft to drive the compressor, restarting the cycle.
  • the combustion can comprises a number of overlapping annular truncated cone sections 46 each having an axially extending slotted rear edge 48 formed integral therewith.
  • the forward portion 59 of each section is formed with circumferential corrugations 52 providing axially extending alternatingly connected ridges 54 and grooves 5'6.
  • the grooves 56 are substantially pointed, dimple-like indentations with a downstream pointing tip.
  • Each of the portions 50 overlaps the rear axial edge 4-3 of the adjacent section in a manner to abut or contact the inner surface of the grooves 56 with the rear edge as shown at 58 in FIGURE 6.
  • the two edges are then joined rigidly to each other by spot welding as indicated at 66.
  • each of the fluid passages has a restricted throat 66 defined by the rear edge 48 of the adjacent sections and the radial taper of portion 50 to cause an increase in the velocity of the air flow through the passage and a squeezing action or diffusion of the air in this region by a reduction in height of the volume of air.
  • the passages are each further tapered circumferentially with respect to each other, the taper 68 diverging laterally with an increase in the axial downstream direction.
  • the quantity or volume of air entering each of the fluid passages is therefore forced out laterally or sideways as the height of the passage is diminished by the longitudinal taper to fill the area defined by the two tapers and the rear edge .48 and wash completely the walls of the grooves and ridges with cool air along the entire axial length thereof.
  • This tapering construction therefore not only provides diverging or fan-shaped intersecting fluid flow paths to cool the portions of the liner around the welds and immediately downstream thereof thereby preventing a burn out of these portions, but also delivers air into the can to cool the entire combustion liner.
  • the axial length of overlap between edge 48 and portion 50 is preferably greater than the axial length of the corrugations to permit forming of the passages in a manner that absolutely makesithe air expand to till the area behind each corrugation.
  • This overlap length will also, of course, vary as a function of the desired restriction of the throat 66 of the fluid passage, i.e., the throat area as determined by the taper angle of portion 50 and the overlap length of edges 48 and 50 will be that area providing the most efiicient restriction to increase the velocity of the fluid passing through the passage to cause a diffusion thereof spreading out the flow against the Walls of the ridges and grooves to completely wash them with cool air and prevent hot spots behind the welds.
  • each section is slotted at 70 to eliminate distortion of the liner by permitting circumferential expansion or contraction thereof during thermal expansion or contraction of the liner.
  • the slots also aid in forming the axially extending edge 48.
  • the first annular section 72 of the combustion liner has an axial forward edge portion 74 cooperating with a corrugated annular spacer element 76 having secured thereto internally thereof the dome 28.
  • the construction of the domeas shown is known and immaterial to an understanding of the invention. Suflice it to say, however, that the dome may have a number of radially and circumferentially spaced primary air holes 78 for the admission of air therethrough, and a number of circumferentially spaced stepper swirler batfle plates 80 cooperating therewith to diifuse the primary air for better fuel-air mixing.
  • the first section liner is also provided with the opening 42 within which a ferrule 82 is inserted and rigidly secured to the can receiving therein the igniter plug 44.
  • vol ume of air admitted through the fluid passages will be squeezed laterally to wash the walls of the grooves and ridges of the corrugations with cool air, thereby completely cooling this section of the liner and providing a cooling layer of fluid between the flame of the combustion can and the liner to act as an insulator to prevent burn out thereof.
  • this invention provides a combustion can of rigid construction, and with means to cool the same eifectively to prevent hot spots therein and subsequent failure thereof.
  • This invention also provides a combustion can that can be manufactured economically and one that has a long endurance life. While the invention has been illustrated in connection with the combustion section of a gas turbine engine, it will be clear to those skilled in the art to which this invention pertains that many modifications can be made thereto Without departing from the scope of the invention.
  • a combustion liner including a number of tele-' scopically mounted sections having overlapping portions, the Walls of the overlapping portions converging radially towards each other in a downstream direction, and a plurality of cir-cumferentially spaced dimple-like indentations formed in the wall of one portion contacting the wall of the other portion, each of the indentations converging circumferentially in a downstream direction thus defining fluid passages between the walls and indentations converging radially and diverging circumferentially in a downstream direction, the edge of one portion extending farther downstream than the downstream end of said indentations, the shape of the passages thereby causing the fluid passing therethrough to follow a fanshaped path to efiectively cool the sections.
  • a combustion liner including a number of telescopically mounted sections have overlapping portions, the walls of-the overlapping portions converging radially towards each other in a downstream direction, and a plurality of circumferentially spaced dimple-like indentations formed in the wall of one portion contacting the wall of the other portion, each of the indentations converging circumferentially in a downstream direction thus defining fluid passages between the walls and indentations converging radially and diverging circumferentially in a downstream direction, the edge of one portion extending farther downstream than the down stream end of said indentations, the shape of said passages thereby effecting a squeezing action on any fluid passing therethrough causing the fluid to follow a fan-shaped path to effectively cool the sections and the points on said portion walls immediately downstream of the points of contactof said Walls.
  • a combustion liner including a number of telescopically mounted sections having overlapping portions, the walls of the overlapping portions converging radially towards each other in a downstream direction, and a plurality of circumferentially spaced substantially triangle-shaped indentations formed in the wall of one portion contacting the wall of the other portions, each of the indentations converging circumferentially in a downstream direction thus defining fluid passages between the walls and indentations that converge radially and diverge circumferentially in a downstream direction, the edge of one portion extending farther downstream than the downstream end of the indentations, the shape of said passages thereby effecting a squeezing action on any fluid passing therethrough causing the fluid to follow a fan-shaped path to effectively cool the sections and the points on said portion walls immediately downstream of the points of contact of said walls.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Spray-Type Burners (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US844279A 1959-10-05 1959-10-05 Combustion liner Expired - Lifetime US3064425A (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
US844279A US3064425A (en) 1959-10-05 1959-10-05 Combustion liner
GB31939/60A GB891836A (en) 1959-10-05 1960-09-16 Improvements in flame tubes for combustion chambers
FR840205A FR1268931A (fr) 1959-10-05 1960-10-03 Tube de flamme pour chambre de combustion
CH1120260A CH380444A (de) 1959-10-05 1960-10-05 Flammrohr für Brennkammern

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US844279A US3064425A (en) 1959-10-05 1959-10-05 Combustion liner

Publications (1)

Publication Number Publication Date
US3064425A true US3064425A (en) 1962-11-20

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Application Number Title Priority Date Filing Date
US844279A Expired - Lifetime US3064425A (en) 1959-10-05 1959-10-05 Combustion liner

Country Status (3)

Country Link
US (1) US3064425A (de)
CH (1) CH380444A (de)
GB (1) GB891836A (de)

Cited By (34)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3485043A (en) * 1968-02-01 1969-12-23 Gen Electric Shingled combustion liner
US3589128A (en) * 1970-02-02 1971-06-29 Avco Corp Cooling arrangement for a reverse flow gas turbine combustor
US3705492A (en) * 1971-01-11 1972-12-12 Gen Motors Corp Regenerative gas turbine system
US3742702A (en) * 1971-01-22 1973-07-03 Gen Motors Corp Regenerative gas turbine system
US3751910A (en) * 1972-02-25 1973-08-14 Gen Motors Corp Combustion liner
US3793827A (en) * 1972-11-02 1974-02-26 Gen Electric Stiffener for combustor liner
US3854285A (en) * 1973-02-26 1974-12-17 Gen Electric Combustor dome assembly
US4008568A (en) * 1976-03-01 1977-02-22 General Motors Corporation Combustor support
US4050241A (en) * 1975-12-22 1977-09-27 General Electric Company Stabilizing dimple for combustion liner cooling slot
US4085580A (en) * 1975-11-29 1978-04-25 Rolls-Royce Limited Combustion chambers for gas turbine engines
US4132066A (en) * 1977-09-23 1979-01-02 United Technologies Corporation Combustor liner for gas turbine engine
US4180972A (en) * 1978-06-08 1980-01-01 General Motors Corporation Combustor support structure
US4244178A (en) * 1978-03-20 1981-01-13 General Motors Corporation Porous laminated combustor structure
US4312186A (en) * 1979-10-17 1982-01-26 General Motors Corporation Shingled laminated porous material
US4315406A (en) * 1979-05-01 1982-02-16 Rolls-Royce Limited Perforate laminated material and combustion chambers made therefrom
US4413477A (en) * 1980-12-29 1983-11-08 General Electric Company Liner assembly for gas turbine combustor
US4887663A (en) * 1988-05-31 1989-12-19 United Technologies Corporation Hot gas duct liner
EP0348500A4 (de) * 1987-12-28 1990-04-10 Sundstrand Corp Ringförmige verbrennungseinheit mit tangentialem kühllufteinspritzen.
US5181377A (en) * 1991-04-16 1993-01-26 General Electric Company Damped combustor cowl structure
WO2004023038A1 (en) * 2002-09-03 2004-03-18 Pratt & Whitney Canada Corp. Stress relief feature for aerated gas turbine fuel injector
US20050056020A1 (en) * 2003-08-26 2005-03-17 Honeywell International Inc. Tube cooled combustor
US20060219191A1 (en) * 2005-04-04 2006-10-05 United Technologies Corporation Heat transfer enhancement features for a tubular wall combustion chamber
FR2887615A1 (fr) * 2005-06-22 2006-12-29 Snecma Moteurs Sa Carenage circulaire de chambre de combustion de turbomachine
FR2896575A1 (fr) * 2006-01-26 2007-07-27 Snecma Sa Chambre de combustion annulaire d'une turbomachine
US20070180827A1 (en) * 2006-02-09 2007-08-09 Siemens Power Generation, Inc. Gas turbine engine transitions comprising closed cooled transition cooling channels
RU2319075C1 (ru) * 2006-05-10 2008-03-10 Открытое акционерное общество "Климов" Жаровая труба камеры сгорания
FR2910597A1 (fr) * 2006-12-22 2008-06-27 Snecma Sa Carenage pour fond de chambre de combustion
US20080155988A1 (en) * 2006-08-28 2008-07-03 Snecma Annular combustion chamber for a turbomachine
RU2343355C2 (ru) * 2006-11-30 2009-01-10 Открытое акционерное общество "Климов" Жаровая труба камеры сгорания газотурбинного двигателя
US20100205969A1 (en) * 2007-10-24 2010-08-19 Man Turbo Ag Burner for a Turbo Machine, Baffle plate for Such a Burner and a Turbo Machine Having Such a Burner
RU205407U1 (ru) * 2020-12-08 2021-07-13 Акционерное общество "Объединенная двигателестроительная корпорация" (АО "ОДК") Жаровая труба с компенсационными щелями
US11085642B2 (en) * 2016-05-23 2021-08-10 Mitsubishi Power, Ltd. Combustor with radially varying leading end portion of basket and gas turbine
RU2805719C1 (ru) * 2023-04-10 2023-10-23 Общество с ограниченной ответственностью Научно-производственное объединение "Базовое машиностроение" Жаровая труба камеры сгорания газотурбинного двигателя ДН80 и ДУ80
US11994291B2 (en) 2022-07-21 2024-05-28 General Electric Company Performance factor for a combustion liner

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2645081A (en) * 1949-08-19 1953-07-14 A V Roe Canada Ltd Spacing means for the wall sections of flame tubes
US2670601A (en) * 1950-10-17 1954-03-02 A V Roe Canada Ltd Spacing means for wall sections of flame tubes
US2884759A (en) * 1956-04-25 1959-05-05 Curtiss Wright Corp Combustion chamber construction

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2645081A (en) * 1949-08-19 1953-07-14 A V Roe Canada Ltd Spacing means for the wall sections of flame tubes
US2670601A (en) * 1950-10-17 1954-03-02 A V Roe Canada Ltd Spacing means for wall sections of flame tubes
US2884759A (en) * 1956-04-25 1959-05-05 Curtiss Wright Corp Combustion chamber construction

Cited By (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3485043A (en) * 1968-02-01 1969-12-23 Gen Electric Shingled combustion liner
US3589128A (en) * 1970-02-02 1971-06-29 Avco Corp Cooling arrangement for a reverse flow gas turbine combustor
US3705492A (en) * 1971-01-11 1972-12-12 Gen Motors Corp Regenerative gas turbine system
US3742702A (en) * 1971-01-22 1973-07-03 Gen Motors Corp Regenerative gas turbine system
US3751910A (en) * 1972-02-25 1973-08-14 Gen Motors Corp Combustion liner
US3793827A (en) * 1972-11-02 1974-02-26 Gen Electric Stiffener for combustor liner
US3854285A (en) * 1973-02-26 1974-12-17 Gen Electric Combustor dome assembly
US4085580A (en) * 1975-11-29 1978-04-25 Rolls-Royce Limited Combustion chambers for gas turbine engines
US4050241A (en) * 1975-12-22 1977-09-27 General Electric Company Stabilizing dimple for combustion liner cooling slot
US4008568A (en) * 1976-03-01 1977-02-22 General Motors Corporation Combustor support
US4132066A (en) * 1977-09-23 1979-01-02 United Technologies Corporation Combustor liner for gas turbine engine
US4244178A (en) * 1978-03-20 1981-01-13 General Motors Corporation Porous laminated combustor structure
US4180972A (en) * 1978-06-08 1980-01-01 General Motors Corporation Combustor support structure
US4315406A (en) * 1979-05-01 1982-02-16 Rolls-Royce Limited Perforate laminated material and combustion chambers made therefrom
US4312186A (en) * 1979-10-17 1982-01-26 General Motors Corporation Shingled laminated porous material
US4413477A (en) * 1980-12-29 1983-11-08 General Electric Company Liner assembly for gas turbine combustor
EP0348500A4 (de) * 1987-12-28 1990-04-10 Sundstrand Corp Ringförmige verbrennungseinheit mit tangentialem kühllufteinspritzen.
US4887663A (en) * 1988-05-31 1989-12-19 United Technologies Corporation Hot gas duct liner
US5181377A (en) * 1991-04-16 1993-01-26 General Electric Company Damped combustor cowl structure
US6823677B2 (en) 2002-09-03 2004-11-30 Pratt & Whitney Canada Corp. Stress relief feature for aerated gas turbine fuel injector
WO2004023038A1 (en) * 2002-09-03 2004-03-18 Pratt & Whitney Canada Corp. Stress relief feature for aerated gas turbine fuel injector
US20050056020A1 (en) * 2003-08-26 2005-03-17 Honeywell International Inc. Tube cooled combustor
US7043921B2 (en) * 2003-08-26 2006-05-16 Honeywell International, Inc. Tube cooled combustor
US7464537B2 (en) * 2005-04-04 2008-12-16 United Technologies Corporation Heat transfer enhancement features for a tubular wall combustion chamber
US20060219191A1 (en) * 2005-04-04 2006-10-05 United Technologies Corporation Heat transfer enhancement features for a tubular wall combustion chamber
FR2887615A1 (fr) * 2005-06-22 2006-12-29 Snecma Moteurs Sa Carenage circulaire de chambre de combustion de turbomachine
FR2896575A1 (fr) * 2006-01-26 2007-07-27 Snecma Sa Chambre de combustion annulaire d'une turbomachine
US20070180827A1 (en) * 2006-02-09 2007-08-09 Siemens Power Generation, Inc. Gas turbine engine transitions comprising closed cooled transition cooling channels
US7827801B2 (en) 2006-02-09 2010-11-09 Siemens Energy, Inc. Gas turbine engine transitions comprising closed cooled transition cooling channels
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
US20080155988A1 (en) * 2006-08-28 2008-07-03 Snecma Annular combustion chamber for a turbomachine
US8387395B2 (en) * 2006-08-28 2013-03-05 Snecma Annular combustion chamber for a turbomachine
RU2343355C2 (ru) * 2006-11-30 2009-01-10 Открытое акционерное общество "Климов" Жаровая труба камеры сгорания газотурбинного двигателя
FR2910597A1 (fr) * 2006-12-22 2008-06-27 Snecma Sa Carenage pour fond de chambre de combustion
US20100205969A1 (en) * 2007-10-24 2010-08-19 Man Turbo Ag Burner for a Turbo Machine, Baffle plate for Such a Burner and a Turbo Machine Having Such a Burner
US11085642B2 (en) * 2016-05-23 2021-08-10 Mitsubishi Power, Ltd. Combustor with radially varying leading end portion of basket and gas turbine
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
RU2805719C1 (ru) * 2023-04-10 2023-10-23 Общество с ограниченной ответственностью Научно-производственное объединение "Базовое машиностроение" Жаровая труба камеры сгорания газотурбинного двигателя ДН80 и ДУ80

Also Published As

Publication number Publication date
GB891836A (en) 1962-03-21
CH380444A (de) 1964-07-31

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