US2676460A - Burner construction of the can-an-nular type having means for distributing airflow to each can - Google Patents

Burner construction of the can-an-nular type having means for distributing airflow to each can Download PDF

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US2676460A
US2676460A US151507A US15150750A US2676460A US 2676460 A US2676460 A US 2676460A US 151507 A US151507 A US 151507A US 15150750 A US15150750 A US 15150750A US 2676460 A US2676460 A US 2676460A
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air
shield
combustion
shields
vanes
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US151507A
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Edmund D Brown
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Raytheon Technologies Corp
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United Technologies Corp
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/46Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings

Description

April 27, 1954 w 2,676,460

BURNER CONSTRUCTION OF T CAN STRIBUTING AIR -ANNULAR TYPE HAVING MEANS FOR DI Filed March 25, 1950 FLOW TO EACH CAN 3 Sheets-Sheet 1 175 .1.

futenior Ed at a? Bwown MW,

Apnl 27, 1954 E. D. BROWN 2,676,460

BURNER CONSTRUCTION OFYTHE CAN-ANNULAR TYPE HAVING MEANS FOR DIS'EIBUTING AIRFLOW TO EACH CAN Filed March 23, 1950 3 Sheets-Sheet 2 April 27, 1954 ROWN 76,460

E. D. B 2 BURNER CONSTRUCTION OF THE CAN-ANNULAR TYPE HAVING MEANS FOR DISTRIBUTING AIRFLOW TO EACH CAN Filed March 25, 1950 3 Sheets-Sheet 3 .and air to mix at a point close to Patented Apr. 27, 1954 NULAR TYPE HAVIN TRIBUTING AIRFLO Edmund D. Brown,

to United Aircraft OF THE CAN -AN G MEANS FOR DIS- W TO EACH CAN Manchester, Cnn., assignor Corporation, East Hartford,

Conn., a corporation of Delaware Application March 23, 1950, Serial No. 151,507

4 Claims. (Cl. 6039.65)

This invention relates to combustion chambers especially adapted for gas turbine power plants.

The copending application of Highberg, Serial No. 150,973 filed March 21, 1950, discloses a type of combustion chamber in which the normally small length-to-width ratio is increased to a more efiicient ratio by providing a plurality of parallelly positioned combustion spaces arranged in side-by-side relation in the combustion chamber such that each space has a relatively large length-to-width ratio. The parallelly arranged combustion spaces provide for a flow of primary and secondary air adjacent to the walls of the combustion chamber externally of the spaces and also between the spaces susbtantially centrally of the gas passage. A feature of the present invention is an arrangement for dividing the air flow between the outer passages adjacent the walls of the duct and the centrally located air passage. In this way the air fiow is divided upstream of the combustion spaces in such a manner as to assure best combustion within the spaces.

The fuel and air are preferably mixed and burned before the mixture reaches the turbine thereby preventing excess temperatures from occurring within the turbine. A feature of the invention is an arrangement for causing the fuel the upstream end of the combustion space in order that combustion will be completed within the limits of the space. Another feature is an arrangement which will tend to produce a reverse flowpof the primary air adjacent the upstream end of the combustion space thereby more efiectively mixing the fuel and air and improving the combustion .characteristics.

low the inside surface of the cup for additionally improving the combustion characteristics particularly in connection with the reverse flow of air above referred to.

One feature is the arrangement of the inlet air openings in the shields which define combustion spaces for suitably mixing the air with the fuel in the proper portions for best combustion and -subsequently mixing the products of combustion with the secondary air flowing through the combustion, chamber around and centrally of the combustion spaces.

Other objects and advantages will be apparent from the specification and claims, and from the accompanying drawings which illustrate an embodiment of the invention.

Fig. 1 is a longitudinal sectional view through a part of a gas turbine power plant showing the combustion chamber.

Fig. 2 is an enlarged view of a portion of the combustion chamber on substantially the same plane.

Fig. 3 is a transverse sectional view through the combustion chamber substantially along the line 3-3 of Fig. 1.

Fig. 4 is a greatly enlarged perspective view of the inlet end of a part of a combustion chamber.

Fig. 5 is a fragmentary sectional view on line 5-5 of Fig. 3.

The invention is shown in connection with a gas turbine power plant in which the compressor 2, the last stage only of which is shown, delivers air under pressure to an annular duct 4 in which fuel is mixed with the air and burned to produce power gases which are then discharged through a turbine nozzle 6 for driving the turbine rotor 8. This rotor, only a part of which is shown, is connected to the compressor rotor ill by a sleeve l2 which may be supported in spaced bearings I 4 and l 6 within the supporting structure IS.

The compressor casing 2b which carries stationary vanes 22 has attached thereto at its downstream end the diifuser section 24 of the combustion chamber. This diffuser section provides an annular passage from the compressor casing and consists in diverging inner and outer Walls 26 and 28 which, adjacent their upstream ends are spaced apart and parallel and are interconnected by radially extending members (to which function as straightening vanes. These walls diverge in a downstream direction and form with the compressor casing 26 a part of a load carrying structure of the power plant. The supporting structure [8 for the bearings l4 and I6 is carried by radially extending members 30 welded or otherwise attached to the diifuser section.

The central part of the combustion chamber downstream of the diffuser section is enclosed within an outer cylindrical wall 36. The upstream end of the wall 34 is bolted to an extension of outer wall 28 of the diffuser section as by bolts 40. The downstream end of the wall 34 has a cylindrical extension 42 somewhat smaller than wall 34 and surrounding the turbine.

Within the wall 34 and supported thereby in closely spaced relation is an annular heat shield 45 which, in effect, forms a continuation of the outer wall of the duct 4. The shield 46 becomes smaller in diameter at its downstream end where it joins a second shield 48 extending diagonally inwardly to engage withthe outer wall of the turbine nozzle 6, this shield 48 forming inefiect a further continuation of the outer wall of the duct 4.

From the downstream end of the inner wall 25 of the diffuser d, the upstream end of which is connected as by bolts to the end of wall 26. 'The'sleeve becomes slightly smaller at it-s downstream'endbut is nearly parallel to the wall :34. The-sleeve 54 is attached at its lower end to a frustoaoonical casing member 55 which becomes gradually larger in diameter to cause inner'wall of duct 4 to converge toward the outer wall 34. 'IIIE'IdOWIL. stream end of member 55 engages with the inner wall of the .turbinenozzle 5. Sleeve 54is shielded from-the heat within the combustion chamber by a liner'55 closelyspacedrfromand on theloutsideof the sleeve 54 and member 55 is shielded by a shield 51. The duct 4 thus-includes the diffuser section in which the inner and outer walls 5diverge, the substantially cylindrical section in which the combustion takesplace, and the section of gradually decreasing cross sectional area 'for guiding the hot power fluid to the turbine nozzle.

The invention is shown in conjunction with a can-annular combustion chamber construction in which'the duct 4 is annular and concentric to the axis of the compressor and turbine while the individual burnercans 58 are arranged in a ring around the axis of the turbine and parallel thereto, being spaced apart angularly within the annular duct. Each burner canincludes an outer substantially cylindrical shield 50 thediameter of which is slightly less than thelradial dimension of the annular duct to provide an airspace 52 between the shield 6 8- and the .annulanheat shield 46 and also to provide an air passage 64between the shield 50 and the liner 56.

Mounted within the cylindrical-shield '50 and substantially concentric-theretoisaninner shield 66 which vprovides a centralairpassage 6.8 the downstream end-of which is .closedat a. point upstream of the'turbinenozzle.

The inner andouter shields .filliand 56 arerhel'd in spaced relation to ieachlotherbyan annular .cap Iii which closes the upstreameendo'f the space between the shields and which supports .a plurality of cups 12 with their-basesfacing upstream and with theopen ends of the .cups coinciding with openings -14 in'thelcap 10. .A fuel nozale15 .is supportedcentrallywithinthe base of each-cup and the nozzle is surrounded by shrouded swirl vanes 18 which are adapted to impartatangential swirl to the air passing over the vanes .an'dbetween the-nozzle and shroud '19 into'thelcup, to cause it,-by centrifugal force, tofollow the-inside surface of the cup and also to producela cylinder of moving air as-itjprogresses .down the stream. The cap Ill carries a number ofthese 'fuel'in'jection cups each of which supports a fuel nozzle with the result that the/flame tube which is the annular space between the shields (inland Q66.is provided with fuel at uniformly spaced points around its circumference thereby assuring -a fairly uniform distribution of the combustion withinthe combustion space.

Only a small amount-of air enters the flame tube past the swirl wanes "l8, thegreatergpartiof thegprimary: air entering the tube throughopenings 80 in the outer shield 261i :andcsimilar repensection there extends a sleeve 4 ings 82 in the inner shields 5B. In the arrangement shown the openings 80 are arranged in spaced parallel rows and these openings may be in direct radial alignment with openings 82 in the inner shields. The .air enters the flame tube through the openings 80 and'82 in arsubstantially radial direction but because of the reduced pressure within the swirling air discharged from the fuel injection cups 12 the primary air entering at least the first row of Inlet openings tends to flow in a substantially-upstream direction against the fuel being injected by nozzles 15 with the result .of improved-mixing and rapid combustion. T0- "ward the downstream end of the shield 60 it may lbe'ladvantageous toj provide elongated slots 84 for the rapid admission of secondary air to mix with -'the'products ofcombustion. It will be apparent that theieffect of the centrally located shields 66 is to reduce the effective width of the combus- 94 extending forwardly tion chamber, since the effective dimension is now the spacing between "the shields 6B and '66, without in any way reducing the length of the flame time so that the resulting :ratio of lengthto-width is increased to amore -effective ratio'for best combustion.

The cups "T2 are in effect large angle diffusers andtheswirl action assists the air "in following the inside walls of the cups,'ther'ebypreventing the 'flow from breaking away from the walls. The diifuser action results in a low pressure directly downstream o'f'the swirl vanes 18 'and 'a higher pressure with lower "velocity :at 'a point approximately in line'withthefiirsti'row "of openings en. This .higher pressure causes a fiow toward the low pressure area, directly 'a'gainstthe fuel sprayed in from then'ozzle. Moreover, the swirling air is provided'by centrifugalQforce with a pressure differential such that "the 'air has "a higher pressureiat'the rim of the cup and a decreasing pressurefrom the'rim toward thetaxis of the cup. Thispressure 'differentialassiststhe air flow entering the openings in penetrating this swirling air'sothat the airfrom the'openings Blimay reach the 'central part of the combustion space.

The circumferential spacing "of the 'openings 80 is such as to 'provide'tapproxima'tly triangular areas of substantial size between the air jets directed into the combustion space from *the openings 80. 'Asa result the air which "had 'in- 'itiallymoved'in anupstream directioniwithinthe combustion space, and the fuel mixedwith the air, and tin which burning 'is'occurring, can "and does when reversed in direction adjacent the open end 'of the cup flow downstream through these areas without interfering with the entry of primary air to the combustion spaces.

Instead of attaching the outer rim 'of "jthecap 10 directly to the outer shield St, 'a 'wiggle "strip spaces'the cap'fromthe'shieldand permitsa flow ofairlongitudinally along the inner surface of the shield for "cooling the shield when combustion is taking place. .A similar wiggle "strip 88 'may be'provi'd'ed between the downstream'end of the shield '61] and the collector ring "'QOwhich, as shown, fits between the converging walls at the'downstream'en'd ofthe combustion chambers adjacent to the turbine. The wiggle strip 88 allows :coolingair to flow overthe inn'er'sur'face of the members!) for cooling purposes.

For the purpose -of assuring the proper distributionof "primary air between the "outer air passages 62 and '54 and the central 'air passage '68, the inlet end of the combustion chamber has a dividing member 92 which has :an inner wall from thezinnerr-shieldiii stream ends as shown and and an outer wall 96 extending in an upstream direction from the outer periphery of the cap 10. The inner and outer walls merge at their upby suitable proportioning of the dimensions of the inner wall 94 and the location of the upstream end of the member 92 with respect to the surrounding diffuser section, it is possible to accurately proportion the air flow so that the air is introduced through the openings 80 and 82 to best advantage. It may be noted that another function of the walls as and 96 is to enclose the fuel manifolding 98 for the nozzle (Band to streamline the air flow into the burner.

The inner wall 94 may be made up of separate elements 94a and 94b with the downstream end of the element 94a slightly larger than the adjacent end of the element 94b to provide a metering orifice 99 which controls the amount of air entering the space between walls 94 and 96 and thus controls the amount of air delivered through the swirl vanes 18 into the cups (2.

Each of the individual cans is supplied with fuel from a surrounding manifold structure I00 located externally of the diffuser section of the burner and having a substantially radially extending inwardly projecting arm H32 provided with fuel passages I04 and IE6 through which fuel is admitted to the manifold 98 above referred to.

In addition to the inner and outer walls 94 and 9B of the dividing member 92, this member also has flanges I08 extending in a circumferential direction with respect to the burner duct, to engage with supporting struts H9 positioned radially between the upstream ends of adjacent burner cans, as shown in Fig. 5. With these flanges in place the desired proportion of air is delivered to the space radially inward of the flanges and also to the space radially outward thereof.

It is to be understood that the invention is not limited to the specific embodiment herein illustrated and described, but may be used in other ways without departure from its spirit as defined by the following claims.

I claim:

1. In a burner construction, an annular duct having an inlet at one end and an outlet at the opposite end for the flow of air therethrough, a plurality of cylindrical shields parallelly arranged within the duct, said shields being spaced from the duct walls and from each other to provide air passages therebetween, a tube-shaped inner shield within each cylindrical shield and defining with each cylindrical shield a substantially annular combustion space of large length-towidth ratio, an annular closure at the upstream end of each combustion space, each closure having a plurality of openings therein, said openings being filled with a fuel nozzle assembly including a downstream facing cup-shaped element, a fuel nozzle within the base of the cup-shaped element and a plurality of vanes surrounding said nozzle and surrounded by said cup-shaped element, said vanes imparting a tangential swirl to the air entering said combustion space through said vanes, said cylindrical shields and said inner shields being perforated to admit air to said combustion spaces, and means upstream of each of said cylindrical shields and closures for distributing airflow to said inner shield, said vanes and around said cylindrical shield.

2. In a burner construction, an annular duct having an inlet at one end and an outlet at the opposite end for the flow of air therethrough, a

vide air passages therebetween tially annular combustion space plurality of cylindrical shields parallelly ar* ranged within the duct, said shields being spaced from the duct walls andfrom each other to pro- ,a tube-shaped inner shield within each cylindrical shield and defining with each cylindrical shield a substanof large lengthto-width ratio, an annular closure at the upstream end of each combustion space, each closure having a, plurality of openings therein, said openings being filled with a fuel nozzle assembly including a downstream facing cup-shaped element, a fuel nozzle within the base of the cupshaped element and a plurality of vanes surrounding said nozzle and surrounded by said cupshaped element, said vanes imparting a tangential swirl to the air entering said combustion space through said vanes, said cylindrical shields and said inner shields being perforated to admit air to said combustion spaces, means for substantially enclosing the upstream end of each of said cylindrical shields, means for admitting air through said enclosing means to said inner shield and said vanes, said enclosing means distributing airflow to said inner shield, said vanes and around said cylindrical shield.

3. In a burner construction, an annular duct having an inlet at one end and an outlet at the opposite end for the flow of air therethrough, a plurality of cylindrical shields parallelly arranged within the duct, said shields being spaced from the duct walls and from each other to provide air passages therebetween, a tube-shaped inner shield within each cylindrical shield and defining with each cylindrical shield a substantially annular combustion space of large lengthto-width ratio, an annular closure at the upstream end of each combustion space, each closure having a plurality of openings therein, said openings being filled with a fuel nozzle assembly including a downstream-facing cup-shaped element, a fuel nozzle within the base of the cupshaped element and a plurality of vanes surrounding said nozzle and surrounded by said cupshaped element, said vanes imparting a tangential swirl to the air entering said combustion space through said vanes, said cylindrical shields and said inner shields being perforated to admit air to said combustion spaces, a conical shield substantially enclosing the upstream end of each of said cylindrical shields, said conical shield having a central opening through which air is admitted to said inner shield and said vanes, said conical shield distributing airflow to said inner shield, said vanes and around said cylindrical shield.

4. In a burner construction, an annular duct having an inlet at one end and an outlet at the opposite end for the flow of air therethrough, a plurality of cylindrical shields parallelly arranged within the duct, said shields being spaced from the duct walls and from each other to provide air passages therebetween, a tube-shaped inner shield within each cylindrical shield and defining with each cylindrical shield a substantially annular combustion space of large length-towidth ratio, an annular closure at the upstream end of each combustion space, each closure having a plurality of openings therein, said openings being filled with a fuel nozzle assembly including a downstream facing cup-shaped element, a fuel nozzle within the base of the cup-shaped element and a plurality of vanes surrounding said nozzle and surrounded by said cup-shaped element, said vanes imparting a tangential swirl to the air entering said combustion space through said vanes,

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Cited By (30)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2832580A (en) * 1955-02-09 1958-04-29 Selas Corp Of America Convection heating unit
US2882681A (en) * 1953-02-24 1959-04-21 Lucas Industries Ltd Air-jacketed annular combustion chambers for jet-propulsion engines, gas turbines or like prime movers
DE1062985B (en) * 1955-12-30 1959-08-06 United Aircraft Corp Combustion chamber for turbojets
DE1118535B (en) * 1959-03-31 1961-11-30 United Aircraft Corp Flame tube for gas turbine combustors
US3018624A (en) * 1954-03-02 1962-01-30 Bristol Siddeley Engines Ltd Flame tubes for use in combustion systems of gas turbine engines
US3046736A (en) * 1958-02-10 1962-07-31 Thompson Ramo Wooldridge Inc Direction control for gelatin monopropellant rocket engine
US3049882A (en) * 1960-05-16 1962-08-21 Gen Electric Combustor construction with means for prevention of hot streaks
US3119234A (en) * 1960-09-13 1964-01-28 Rolls Royce Combustion chamber for a gas turbine engine
DE1171207B (en) * 1960-04-01 1964-05-27 United Aircraft Corp Canned burner for a gas turbine combustor
US4173118A (en) * 1974-08-27 1979-11-06 Mitsubishi Jukogyo Kabushiki Kaisha Fuel combustion apparatus employing staged combustion
EP0169431A1 (en) * 1984-07-10 1986-01-29 Hitachi, Ltd. Gas turbine combustor
US4720970A (en) * 1982-11-05 1988-01-26 The United States Of America As Represented By The Secretary Of The Air Force Sector airflow variable geometry combustor
US5161367A (en) * 1991-04-18 1992-11-10 Westinghouse Electric Corp. Coal fired gas turbine system with integral topping combustor
US5274991A (en) * 1992-03-30 1994-01-04 General Electric Company Dry low NOx multi-nozzle combustion liner cap assembly
US5357745A (en) * 1992-03-30 1994-10-25 General Electric Company Combustor cap assembly for a combustor casing of a gas turbine
US5497613A (en) * 1993-12-03 1996-03-12 Westinghouse Electric Corporation Hot gas manifold system for a dual topping combustor gas turbine system
US6182451B1 (en) * 1994-09-14 2001-02-06 Alliedsignal Inc. Gas turbine combustor waving ceramic combustor cans and an annular metallic combustor
US20050279077A1 (en) * 2004-06-18 2005-12-22 General Electric Company Off-axis pulse detonation configuration for gas turbine engine
US20090223227A1 (en) * 2008-03-05 2009-09-10 General Electric Company Combustion cap with crown mixing holes
US20100186414A1 (en) * 2008-12-15 2010-07-29 Sonic Blue Aerospace, Inc. Magnetic ion plasma annular injection combustor
US20100199684A1 (en) * 2008-12-31 2010-08-12 Edward Claude Rice Combustion liner assembly support
US20110154825A1 (en) * 2009-12-30 2011-06-30 Timothy Carl Roesler Gas turbine engine having dome panel assembly with bifurcated swirler flow
US20110197586A1 (en) * 2010-02-15 2011-08-18 General Electric Company Systems and Methods of Providing High Pressure Air to a Head End of a Combustor
US20120117976A1 (en) * 2010-11-11 2012-05-17 General Electric Company Apparatus and method for igniting a combustor
US20130067932A1 (en) * 2011-09-20 2013-03-21 Honeywell International Inc. Combustion sections of gas turbine engines with convection shield assemblies
JP2013213655A (en) * 2012-04-03 2013-10-17 General Electric Co <Ge> Combustor with non-circular head end
US20130305725A1 (en) * 2012-05-18 2013-11-21 General Electric Company Fuel nozzle cap
US20150040579A1 (en) * 2013-08-06 2015-02-12 General Electric Company System for supporting bundled tube segments within a combustor
JP2015513063A (en) * 2012-03-29 2015-04-30 エクソンモービル アップストリーム リサーチ カンパニー Turbomachine combustor assembly
US9181812B1 (en) * 2009-05-05 2015-11-10 Majed Toqan Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines

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Cited By (39)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2882681A (en) * 1953-02-24 1959-04-21 Lucas Industries Ltd Air-jacketed annular combustion chambers for jet-propulsion engines, gas turbines or like prime movers
US3018624A (en) * 1954-03-02 1962-01-30 Bristol Siddeley Engines Ltd Flame tubes for use in combustion systems of gas turbine engines
US2832580A (en) * 1955-02-09 1958-04-29 Selas Corp Of America Convection heating unit
DE1062985B (en) * 1955-12-30 1959-08-06 United Aircraft Corp Combustion chamber for turbojets
US3046736A (en) * 1958-02-10 1962-07-31 Thompson Ramo Wooldridge Inc Direction control for gelatin monopropellant rocket engine
DE1118535B (en) * 1959-03-31 1961-11-30 United Aircraft Corp Flame tube for gas turbine combustors
DE1171207B (en) * 1960-04-01 1964-05-27 United Aircraft Corp Canned burner for a gas turbine combustor
US3049882A (en) * 1960-05-16 1962-08-21 Gen Electric Combustor construction with means for prevention of hot streaks
US3119234A (en) * 1960-09-13 1964-01-28 Rolls Royce Combustion chamber for a gas turbine engine
DE1217139B (en) * 1960-09-13 1966-05-18 Rolls Royce Combustor for a gas turbine plant
US4173118A (en) * 1974-08-27 1979-11-06 Mitsubishi Jukogyo Kabushiki Kaisha Fuel combustion apparatus employing staged combustion
US4720970A (en) * 1982-11-05 1988-01-26 The United States Of America As Represented By The Secretary Of The Air Force Sector airflow variable geometry combustor
EP0169431A1 (en) * 1984-07-10 1986-01-29 Hitachi, Ltd. Gas turbine combustor
US4898001A (en) * 1984-07-10 1990-02-06 Hitachi, Ltd. Gas turbine combustor
US5161367A (en) * 1991-04-18 1992-11-10 Westinghouse Electric Corp. Coal fired gas turbine system with integral topping combustor
US5274991A (en) * 1992-03-30 1994-01-04 General Electric Company Dry low NOx multi-nozzle combustion liner cap assembly
US5357745A (en) * 1992-03-30 1994-10-25 General Electric Company Combustor cap assembly for a combustor casing of a gas turbine
US5497613A (en) * 1993-12-03 1996-03-12 Westinghouse Electric Corporation Hot gas manifold system for a dual topping combustor gas turbine system
US6182451B1 (en) * 1994-09-14 2001-02-06 Alliedsignal Inc. Gas turbine combustor waving ceramic combustor cans and an annular metallic combustor
US7200987B2 (en) * 2004-06-18 2007-04-10 General Electric Company Off-axis pulse detonation configuration for gas turbine engine
US20050279077A1 (en) * 2004-06-18 2005-12-22 General Electric Company Off-axis pulse detonation configuration for gas turbine engine
US20090223227A1 (en) * 2008-03-05 2009-09-10 General Electric Company Combustion cap with crown mixing holes
US20100186414A1 (en) * 2008-12-15 2010-07-29 Sonic Blue Aerospace, Inc. Magnetic ion plasma annular injection combustor
US9046272B2 (en) * 2008-12-31 2015-06-02 Rolls-Royce Corporation Combustion liner assembly having a mount stake coupled to an upstream support
US20100199684A1 (en) * 2008-12-31 2010-08-12 Edward Claude Rice Combustion liner assembly support
US9181812B1 (en) * 2009-05-05 2015-11-10 Majed Toqan Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines
US9027350B2 (en) * 2009-12-30 2015-05-12 Rolls-Royce Corporation Gas turbine engine having dome panel assembly with bifurcated swirler flow
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