US3851466A - Combustion apparatus - Google Patents

Combustion apparatus Download PDF

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Publication number
US3851466A
US3851466A US00350383A US35038373A US3851466A US 3851466 A US3851466 A US 3851466A US 00350383 A US00350383 A US 00350383A US 35038373 A US35038373 A US 35038373A US 3851466 A US3851466 A US 3851466A
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United States
Prior art keywords
zone
liner
air
combustion
casing
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Expired - Lifetime
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US00350383A
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English (en)
Inventor
A Verdouw
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
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Motors Liquidation Co
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Publication date
Application filed by Motors Liquidation Co filed Critical Motors Liquidation Co
Priority to US00350383A priority Critical patent/US3851466A/en
Priority to CA187,198A priority patent/CA982831A/en
Priority to DE2415036A priority patent/DE2415036C2/de
Priority to GB1499874A priority patent/GB1423799A/en
Priority to JP4009874A priority patent/JPS5421482B2/ja
Application granted granted Critical
Publication of US3851466A publication Critical patent/US3851466A/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
    • 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
    • F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow

Definitions

  • a combustion apparatus for a regenerative gas turbine engine includes a combustion liner providing for prevaporization and premixing of fuel, for extended residence time after primary combustion, and for delayed dilution of the combustion products.
  • the liner includes a premix-prevaporiation zone with relatively small crosssectional area, with fuel injecting means at the upstream end. Air flows without swirl through this zone, and through a flare in the liner where the cross section area increases greatly into the combustion zone.
  • a centerbody in the flare directs flow along the inner surface of the wall and additionally heats the fuel-air mixture.
  • Supports for the centerbody define spaced chutes for the air entering the combustion apparatus, and vanes extending inwardly from the wall downstream of the flare deflect the flow from the chutes toward the center of the combustion zone.
  • Combustion air is admitted only through the premix-prevaporization zone.
  • the combuston zone and part of a residence zone downstream of the combustionzone are cooled by convection only, using dilution air. Entrance holes for dilution air are near the downstream end of the liner.
  • My invention is directed to combustion apparatus, particularly to combustion apparatus for use at relatively high air inlet temperatures, and usually at considerable pressure, as contrasted to operation near normal atmospheric conditions.
  • the preferred embodiment of my invention is intended as a combustion apparatus for a regenerative gas turbine engine such as might be used for vehicle or aircraft propulsion.
  • Such engines operate at pressure ratios of the order of four orfive to one, which means that the pressure of the air in the combustion apparatus under sea level standard conditions rises to 45 to 60 psig.
  • the air is heated to some extent by the compression; in a regenerative engine, it is additionally heated by heat transfer from air exhausted from the engine, so that the air entering the combustion apparatus under full' power conditions may be at about 900 F.
  • Prior art combustion chambers for engines of this sort have incorporated combustion liners within which the combustion takes place, and the principal emphasison design of such liners has been to secure dependable burning of the fuel, minimize pressure drop, minimize the volume of the combustion liner, and provide a durable and trouble-free structure.
  • pollutants are principally particulates (normally unburned carbon), carbon monoxide, unburned hydrocarbons, and nitrogen oxides resulting from the combination of atmospheric nitrogen and oxygen in the hot combustionzone.
  • the preferred combustion apparatus according to my invention involves the following'features to improve emission characteristics:
  • a premix-prevaporization section is provided up'- stream of the primary combustion section to premix the fuel and primary 'air and to'preva'porize the fuel ahead ofthe combustion or reaction zone. This is to improve combustor homogeneity and to avoid fuel droplet burning, thereby eliminating carbon entirely.
  • the liner provides for sudden expansion from the premix-prevaporization zone to the combustion zone for flame stabilization.
  • the primary combustion zone is operated somewhat lean to minimize formation of oxides of nitrogen.
  • Dilution of the combustion products is delayed to provide a relatively extended residence of the combustion products for complete consumption of carbon monoxide and hydrocarbons.
  • the combustion zone is cooled by convection cooling rather than film cooling to avoid quenching of the combustion reactions in relatively cold film cooling air as in prior combustion liners.
  • combustion liner having a relatively great length between the point of initiation of flame and the point of dilution or quenching of the flame; to provide combustion apparatus in which a fuel-air mixture flows over a heated centerbody from a premix-prevaporization zone of low cross-sectional area into a combustion zone of large cross-sectional area; and, in general, to provide combustion apparatus that is better suited to the requirements of practice than those now available.
  • FIG. 1 is a longitudinal sectional view of a combustion apparatus embodying the invention.
  • FIG. 2 is a cross-section view of the same taken on the plane indicated by the line 2-2 in FIG. 1.
  • FIG. 3 is a cross-section view taken on the plane indicated by the line 3-3 in FIG. 1.
  • the drawings represent the liner according to my invention as installed in combustion apparatus similar to that of the well known T63 aircraft engine, in which the combustion apparatus is of a single can type and the single liner discharges-into an annular turbine inlet. In that engine, compressed air issupplied to the combustion apparatus through air tubes entering'the side of a combustion casing. It should be understood, however, that my invention is directed particularly to the structure of the liner and to its cooperation with the combustion outer'casing and that various'types of combustion casings or air supply means may be used. Also, the liner need not have an annular outlet. Also, the principles of the invention may be embodied in an annular combustion liner in which the liner is constituted by outer and inner wallshaving an annular combustion space betweenfthem.
  • the combustion apparatus includes a casing 2 adapted to receive air under pressure and a liner 3.
  • the casing 2 is bolted at flanges 4 to an outer case 6 of the turbine of a gas turbine engine.
  • the casing 2 includes a generally cylindrical portion extending in the upstream direction from the flanges 4 and an enlarged upstream portion 7 closed by an end cover 8.
  • the combustion air is discharged through air tubes 10 into a plenum 11 defined by the upstream portion of the casing.
  • the liner 3 which is of circular cross-section in the form illustrated, has a side wall 12 defining in succession from the upstream end of the liner a primary or combustion air entrance 14, a generally cylindrical premix-prevaporization zone 15, a flare 16, a combustion zone 18, a residence zone 19, and a dilution zone 20.
  • the downstream end of the liner fits within an outer turbine shroud 22, and a flange 23 on the exterior of the downstream end of the liner fits over shroud 22.
  • the liner is thus located at its downstream or discharge end on the turbine outer shroud.
  • An annular baffle 24 mounted on the turbine defines with the shroud 22 an annular entrance 26 into the turbine for the motive fluid generated in the combustion apparatus.
  • a wall indicated schematically at 27 blocks flow of air through the gap between the casing and liner.
  • the upstream end of liner 3 is supported on a fuel injection ring 28 disposed within the entrance 14 and supported by the fuel tubes 30 or other supporting means from a fitting 31 suitably fixed to the end cover 8.
  • Fuel is supplied to the combustion apparatus through a fuel tube 32.
  • the combustion apparatus is designed for use with a liquid hydrocarbon fuel.
  • the fuel injection ring 28 thus supports the upstream end of the liner.
  • the fuel injection ring which is of known type, is considered to be the most suitable means for carburetting the air flowing into the inlet.
  • the fuel injection ring 28 is a composite circular ring defining an internal circular manifold 34 for fuel supplied through tubes 30 and having ports 35 extending from the manifold and directed tangentially to the interior surface of ring 28 to discharge fuel in the form of an annulus on the inner surface of the ring. This fuel is discharged behind a step or shoulder 36 on the interiorof the ring.
  • struts 38 extend from the exterior of ring 28 to the wall of the liner. Air under pressure flows from the plenum 11 over the exterior and through the interior of fuel injection ring 28 and downstream through the premix-prevaporization zone 15. The fuel is airblast atomized off ring 28 by the air flowing past it. During travel through the premix-prevaporization zone, the heated air evaporates the fuel so as to provide a substantially homogeneous mixture of air and vaporized fuel. It should be noted that the struts 38 are not swirlers, and that the air preferably flows through the zone with no circumferential component of velocity.
  • a centerbody 39 of circular cross section the outer surfaceof which roughly parallels the inner surface of the liner wall.
  • the centerbody is'hollow and is open at'its downstream end so that heat from combustion in the combustion zone heats the centerbody, thus contributing a slight amount of heat to the air-fuel mixture flowing over it.
  • the centerbody is supported from the flaring wall portion 16 by eight generally V-shaped sheet metal struts or air guides 40 distributed around the circumference of the centerbody which may be welded to the centerbody and liner wall. These air guides define eight air chutes 42 between the struts. The chute outlets occupy approximately half of the circumference of the annular passage 43 between the centerbody and the liner wall.
  • the centerbody serves to deflect the flow of air and fuel along the inside of the wall 16 and the chutes concentrate this flow into eight jets evenly distributed around the circumference.
  • the centerbody and struts are heated by the combustion just downstream of them and therefore these heated bodies additionally heat the incoming air and fuel somewhat.
  • the centerbody 39 acts as a flow director. creating a turbulent zone behind it to retain the flame against blowing out downstream of the liner. From the flare 16, the wall 12 of the liner continues first cylindrical and then slightly diverging to the downstream end of the liner. A thorough mixing of the fuel and air in the combustion zone 18 is promoted by a ring of eight turning vanes 44 welded to the wall, the form of which will be apparent from FIGS. 2 and 3. These deflect the incoming air-fuel mixture flowing through the chutes 42 toward the axis of the combustion liner and promote formation of a generally toroidal vortex flow and recirculation in the combustion zone to stabilize combustion. The combustion occurs in the region near the flare l6 and vanes 44 and the resulting combustion products, in which the fuel is almost entire burned, flow through the residence zone 19 toward the combustion chamber outlet.
  • approximately 40 percent of the total air supplied to the combustion apparatus enters through the air entrance l4 and constitutes combustion air.
  • the remaining 60 percent is dilution air which flows radially inward from the casing through a ring of six large dilution air holes 46 disposed near the downstream end of the liner.
  • the space abreast of these holes and extending to the outlet of the liner constitutes the dilution zone 20.
  • the proportions of primary air and dilution air may, of course, be varied to suit a particular installation.
  • the dilution air before entering the liner through the holes 46, is used to cool the outer surface of the liner from the flare 16 downstream by convection cooling, avoiding any introduction of air into the liner for film cooling, for instance. Cooling is promoted by an annular shroud 47 surrounding and spaced from the wall 12 of the downstream portion of the liner.
  • the shroud 47 is mounted over; the wall 12 with freedom for relative expansion of the parts by a structure including six plates 48 extending radially outwardly from the inner 7 wall adjacent the flare 16 and six similar plates 50 at the downstream end of the shroud 47. These plates are slidable in slots in the leading and trailing edges of the shroud 47 so the liner may expand radially relative to the cooler shroud.
  • Plates 48 may be welded to the liner before the shroud is put in place andplates then may be inserted in the slots of the shroud and welded to the liner wall.
  • a ring of spacers 51 disposed around the outer surface of the liner wall 12 near the middle of the length of the shroud prevent distortion of this portion of the shroud from partially closing off the annular flow passage 52 between the wall and shroud.
  • a baffle ring or blocking member 54 fixed to the outer surface of shroud 47 extends across the gap between the shroud and the case 2 to minimize or control flow over the outside of the shroud by-passing the cooling air path 52. This blocking member may be considered to divide the space within the casing into an upstream portion ahead of the blocking member and a downstream portion surrounding the residence and dilution zones.
  • Hot air under compression ordinarily on the order of four to five atmospheres and heated at about 900 F. in the regenerator (at full power) is introduced into the plenum l1 and hydrocarbon fuel is introduced through pipe 32 into the annular fuel injection ring 28.
  • the fuel is discharged circumferentially around the inand then are quenched and diluted by the large quantity of air entering through dilution holes 46.
  • the resulting mixture which is of a temperature suitable for use in the turbine, is discharged through the annular outlet 26 into the turbine.
  • the equivalence ratio in the primary section of the specific liner described is calculated to be 0.73 under full load conditions. This varies with load conditions, and various fullload values may be adopted. There are reasons to believe a lower equivalence ratio of the order of 0.5 at full load would provide better emission characteristics.
  • the advantages of my combustion apparatus are greatest where the incoming air is relatively hot. This is the case with normal regenerative gas turbine engines. It also may be the case with very high pressure ratio engines in which the work of compression raises the incoming air to relatively high temperatures. Also, in some cases, the combustion apparatus may work with air preheated otherwise than by a regenerator. In any event, the combustion chamber is particularly effective with air in the temperature range of 700 to 900 F. or above.
  • the apparatus described is well adapted to employ the modes of reduction of undesired contaminants outlined in the introduction to this specification. It will also be seen that it is a relatively simple structure and, while larger than prior art combustion apparatus for such purposes, is not unduly bulky. Specifically, the liner, which is shown to scale in the drawing, is 15 inches long and 6% inches in diameter at its outlet.- it is about 6 inches longer than a typical prior art liner of similar capacity.
  • a combustion apparatus for a regenerative gas turbine engine or the like comprising, in combination, a casing having an upstream portion and a downstream portion and having an entrance for hot compressed air in the upstream portion; a combustion liner disposed within the casing having an upstream end in the upstream portion of the casing arid a downstream end in the downstream portion of the casing; the casing and liner defining between them a space for the compressed air discharging into the liner, and the liner defining an outlet for combustion products at its downstream end; the liner including wall means defining, in flow sequence from its upstream end, a premix-prevaporization zone, a combustion zone, and a dilution zone; the
  • premix-prevaporization zone having an open upstream end defining a primary air inlet adapted to admit air flowing axially of the zone without significant circumferential velocity; means in the air inlet for carburetting the air; the wall means defining a flare from the premix-prevapo'rization zone into the combustion zone; a centerbody disposed centrally of the flare effective to guide the flow along the inner surface of the wall means at the flare, the centerbody serving to heat additionally the carburetted air; the combustion zone being located immediately downstream'of the flare and having subtrance holes through the wall means adjacent to the downstream end of the liner to admit dilution air from the downstream portion of the casing; and the liner wall means between the premix-preva'porization and dilution zones being substantially imperforate.
  • a combustion apparatus for-a regenerative gas turbine engine or the like comprising, in combination, a casing having an upstream portion and a downstream portion and having an entrance for hot compressed air in the upstream portion; a combustion liner disposed within the casing having an upstream end in the upstream portion of the casing and a downstream end in the downstream portion of the casing; the casing and liner defining between them a space for the compressed air discharging into the liner, and the liner defining an outlet for combustion products at its downstream end;
  • the liner including wall means defining, in flow sequence from its upstream end, a premix-prevaporization zone, a combustion zone, and a dilution zone; the premix-prevaporization zone having an open upstream end defining a primary air inlet adapted to admit air flowing axially of the zone without significant circumferential velocity; means in the air inlet for carburetting the air; the wall means defining a flare from the premix-prevaporization zone into the combustion zone; a centerbody disposed centrally of the flare effective to guide the flow along the inner surface of the wall means at the flare, the centerbody serving to heat additionally the carburetted air; guide means connecting the centerbody to the wall means defining spaced chutes for the air and blocking flow between the chutes; the combustion zone being located immediately downstream of the flare and having substantially greater cross-sectional area than the premix-prevaporization zone, and including means effective to promote circulation and mixing in the combustion zone; shroud means spaced outwardly from the wall means extending downstream from the flare over the combustion zone effective
  • a combustion apparatus for a regenerative gas turbine engine or the like comprising, in combination, a casing having an upstream portion and a downstream portion and having an entrance for hot compressed air in the upstream portion; a combustion liner disposed within the casing having an upstream end in the upstream portion of the casing and a downstream end in the downstream portion of the casing; the casing and liner defining between them a space for the compressed air discharging into the liner, and the liner defining an outlet for combustion products at its downstream end; the liner including wall means defining, in flow sequence from its upstream end, a premix-prevaporization zone, a combustion zone, and a dilution zone; the premix-prevaporization zone having an open upstream end defining a primary air inlet adapted to admit air flowing axially of the zone without significant circumferential velocity; means in the air inlet for carburetting the air; the wall means defining a flare from the premix-prevaporization zone into the combustion zone; a centerbody disposed central
  • a combustion apparatus for a regenerative gas turbine engine or the like comprising, in combination, a casing having an upstream portion and a downstream portion and having an entrance for hot compressed air in the upstream portion; a combustion liner disposed within the casing having an upstream end in the upstream portion of the casing and a downstream end in the downstream portion of the casing; the casing and liner defining between them a space for the compressed air discharging into the liner, and the liner defining an outlet for combustion products at its downstream end; the liner including wall means defining, in flow sequence from its upstream end, a premix-prevaporization zone, a combustion zone, a residence zone, and a dilution zone; the premix-prevaporization zone having an open upstream end defining a primary air inlet adapted to admit air flowing axially of the zone without significant circumferential velocity; means in the air inlet for carburetting the air; the wall means defining a flare from the premix-prevaporization zone into the combustion zone; a
  • a combustion apparatus for a regenerative gas turbine engine or the like comprising, in combination, a casing having an upstream portion and a downstream portion and having an entrance for hot compressed air in the upstream portion; a combustion liner disposed within the casing having an upstream end in the upstream portion of the casing and a downstream end in the downstream portion of the casing; the casing and liner defining between them a space for the compressed air discharging into the liner, and the liner defining an outlet for combustion products at its downstream end;
  • the liner including wall means defining, in flow sequence from its upstream end, a premix-prevaporization zone, a combustion zone, a residence zone, and a dilution zone; the premix-prevaporization zone having an open'upstream end defining a primary air inlet adapted to admit air flowing axially of the zone without significant circumferential velocity; means in the air inlet for carburetting the air; the wall means defining a flare from the premix-prevaporization zone into the combustion zone; a centerbody disposed centrally of the flare effective to guide the flow along the inner surface of the wall means at the flare, the centerbody serving to heat additionally the carburetted air; guide means connecting the centerbody to the wall means defining spaced chutes for the air and blocking flow between the chutes; the combustion zone being located immediately downstream of the flare and having substantially greater cross-sectional area than the premix-prevaporization zone, and including stirring vanes extending from the wall means into the path of flow from the chutes effective to promote circulation and mixing in the combustion zone; sh

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Of Fluid Fuel (AREA)
  • Air Supply (AREA)
US00350383A 1973-04-12 1973-04-12 Combustion apparatus Expired - Lifetime US3851466A (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US00350383A US3851466A (en) 1973-04-12 1973-04-12 Combustion apparatus
CA187,198A CA982831A (en) 1973-04-12 1973-12-03 Combustion apparatus for an engine with hot combustion air
DE2415036A DE2415036C2 (de) 1973-04-12 1974-03-26 Brennkammer für Gasturbinentriebwerke mit Regenerativ-Wärmetauschern
GB1499874A GB1423799A (en) 1973-04-12 1974-04-04 Combustion apparatus
JP4009874A JPS5421482B2 (de) 1973-04-12 1974-04-10

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Application Number Priority Date Filing Date Title
US00350383A US3851466A (en) 1973-04-12 1973-04-12 Combustion apparatus

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US3851466A true US3851466A (en) 1974-12-03

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US (1) US3851466A (de)
JP (1) JPS5421482B2 (de)
CA (1) CA982831A (de)
DE (1) DE2415036C2 (de)
GB (1) GB1423799A (de)

Cited By (29)

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Publication number Priority date Publication date Assignee Title
US3982392A (en) * 1974-09-03 1976-09-28 General Motors Corporation Combustion apparatus
US4012904A (en) * 1975-07-17 1977-03-22 Chrysler Corporation Gas turbine burner
US4078377A (en) * 1974-01-28 1978-03-14 Ford Motor Company Internally vaporizing low emission combustor
US4081958A (en) * 1973-11-01 1978-04-04 The Garrett Corporation Low nitric oxide emission combustion system for gas turbines
US4084371A (en) * 1974-07-24 1978-04-18 Howald Werner E Combustion apparatus including an air-fuel premixing chamber
US4113425A (en) * 1975-05-30 1978-09-12 Caloric Gesellschaft Fuer Apparatebau M.B.H Burner for fluid fuels
US4173118A (en) * 1974-08-27 1979-11-06 Mitsubishi Jukogyo Kabushiki Kaisha Fuel combustion apparatus employing staged combustion
US4412414A (en) * 1980-09-22 1983-11-01 General Motors Corporation Heavy fuel combustor
US4483138A (en) * 1981-11-07 1984-11-20 Rolls-Royce Limited Gas fuel injector for wide range of calorific values
US4563875A (en) * 1974-07-24 1986-01-14 Howald Werner E Combustion apparatus including an air-fuel premixing chamber
EP0478305A2 (de) * 1990-09-26 1992-04-01 Hitachi, Ltd. Brennkammer und Verbrennungsvorrichtung
US5117636A (en) * 1990-02-05 1992-06-02 General Electric Company Low nox emission in gas turbine system
US5285631A (en) * 1990-02-05 1994-02-15 General Electric Company Low NOx emission in gas turbine system
US5328355A (en) * 1991-09-26 1994-07-12 Hitachi, Ltd. Combustor and combustion apparatus
US5592819A (en) * 1994-03-10 1997-01-14 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Pre-mixing injection system for a turbojet engine
US5765363A (en) * 1993-07-07 1998-06-16 Mowill; R. Jan Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission
US5924276A (en) * 1996-07-17 1999-07-20 Mowill; R. Jan Premixer with dilution air bypass valve assembly
US6220034B1 (en) 1993-07-07 2001-04-24 R. Jan Mowill Convectively cooled, single stage, fully premixed controllable fuel/air combustor
WO2001040713A1 (en) * 1999-12-03 2001-06-07 Mowill Rolf Jan Cooled premixer exit nozzle for gas turbine combustor and method of operation therefor
US6250066B1 (en) 1996-11-26 2001-06-26 Honeywell International Inc. Combustor with dilution bypass system and venturi jet deflector
WO2002088602A1 (en) * 2001-04-25 2002-11-07 Pratt & Whitney Canada Corp. Turbine premixing combustor
EP1375891A1 (de) * 1997-07-15 2004-01-02 New Power Concepts LLC Sammelrohr für eine Stirlingmaschine
US6925809B2 (en) 1999-02-26 2005-08-09 R. Jan Mowill Gas turbine engine fuel/air premixers with variable geometry exit and method for controlling exit velocities
US20060035183A1 (en) * 2003-02-14 2006-02-16 Richard Carroni Mixer
US20060107647A1 (en) * 2004-04-20 2006-05-25 Labala Gustavo F Turbine, particularly useful for small aircraft
US20070130951A1 (en) * 2005-12-10 2007-06-14 Seoul National University Industry Foundation Combustor
US7513098B2 (en) 2005-06-29 2009-04-07 Siemens Energy, Inc. Swirler assembly and combinations of same in gas turbine engine combustors
EP2436977A1 (de) * 2010-09-30 2012-04-04 Siemens Aktiengesellschaft Brenner für eine Gasturbine
US20140102572A1 (en) * 2012-10-17 2014-04-17 Delavan Inc. Radial vane inner air swirlers

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JPS5232410A (en) * 1975-09-09 1977-03-11 Nissan Motor Co Ltd Combustor for gas turbine engine
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JPS5847610B2 (ja) * 1980-09-29 1983-10-24 株式会社日立製作所 ガスタ−ビン燃焼器
GB2123946B (en) * 1982-05-24 1986-05-21 Freiberg Brennstoffinst Flat flame burner
DE19509854C2 (de) * 1994-09-28 2001-01-04 Abig Werke Carry Gross Gmbh Heizvorrichtung zum Verbrennen von zugeführtem Brennstoff
DE19547913A1 (de) * 1995-12-21 1997-06-26 Abb Research Ltd Brenner für einen Wärmeerzeuger
DE19547912A1 (de) * 1995-12-21 1997-06-26 Abb Research Ltd Brenner für einen Wärmeerzeuger
JP6494563B2 (ja) * 2016-05-13 2019-04-03 ユニ・チャーム株式会社 ペット用介護用品
GB2565761A (en) * 2017-07-28 2019-02-27 Tunley Enginering Combustion engine fuel mixture system

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US2424765A (en) * 1942-10-06 1947-07-29 Stewart Warner Corp Hot-air heater having means to recirculate cooled gases
US2614384A (en) * 1945-01-16 1952-10-21 Power Jets Res & Dev Ltd Gas turbine plant having a plurality of flame tubes and axially slidable means to expose same
GB791617A (en) * 1953-12-11 1958-03-05 Rolls Royce Improvements in or relating to combustion equipment for gas-turbine engines
AT239609B (de) * 1962-08-02 1965-04-12 Prvni Brnenska Strojirna Zd Y Brennkammer-Gasluftmischer
GB1259124A (de) * 1968-12-06 1972-01-05
US3691762A (en) * 1970-12-04 1972-09-19 Caterpillar Tractor Co Carbureted reactor combustion system for gas turbine engine

Cited By (34)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4081958A (en) * 1973-11-01 1978-04-04 The Garrett Corporation Low nitric oxide emission combustion system for gas turbines
US4078377A (en) * 1974-01-28 1978-03-14 Ford Motor Company Internally vaporizing low emission combustor
US4563875A (en) * 1974-07-24 1986-01-14 Howald Werner E Combustion apparatus including an air-fuel premixing chamber
US4084371A (en) * 1974-07-24 1978-04-18 Howald Werner E Combustion apparatus including an air-fuel premixing chamber
US4173118A (en) * 1974-08-27 1979-11-06 Mitsubishi Jukogyo Kabushiki Kaisha Fuel combustion apparatus employing staged combustion
US3982392A (en) * 1974-09-03 1976-09-28 General Motors Corporation Combustion apparatus
US4113425A (en) * 1975-05-30 1978-09-12 Caloric Gesellschaft Fuer Apparatebau M.B.H Burner for fluid fuels
US4012904A (en) * 1975-07-17 1977-03-22 Chrysler Corporation Gas turbine burner
US4412414A (en) * 1980-09-22 1983-11-01 General Motors Corporation Heavy fuel combustor
US4483138A (en) * 1981-11-07 1984-11-20 Rolls-Royce Limited Gas fuel injector for wide range of calorific values
US5117636A (en) * 1990-02-05 1992-06-02 General Electric Company Low nox emission in gas turbine system
US5285631A (en) * 1990-02-05 1994-02-15 General Electric Company Low NOx emission in gas turbine system
EP0478305A2 (de) * 1990-09-26 1992-04-01 Hitachi, Ltd. Brennkammer und Verbrennungsvorrichtung
EP0478305A3 (en) * 1990-09-26 1993-11-24 Hitachi Ltd Combustor and combustion apparatus
US5328355A (en) * 1991-09-26 1994-07-12 Hitachi, Ltd. Combustor and combustion apparatus
US5765363A (en) * 1993-07-07 1998-06-16 Mowill; R. Jan Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission
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US5592819A (en) * 1994-03-10 1997-01-14 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Pre-mixing injection system for a turbojet engine
US5924276A (en) * 1996-07-17 1999-07-20 Mowill; R. Jan Premixer with dilution air bypass valve assembly
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EP1375891A1 (de) * 1997-07-15 2004-01-02 New Power Concepts LLC Sammelrohr für eine Stirlingmaschine
US6925809B2 (en) 1999-02-26 2005-08-09 R. Jan Mowill Gas turbine engine fuel/air premixers with variable geometry exit and method for controlling exit velocities
WO2001040713A1 (en) * 1999-12-03 2001-06-07 Mowill Rolf Jan Cooled premixer exit nozzle for gas turbine combustor and method of operation therefor
WO2002088602A1 (en) * 2001-04-25 2002-11-07 Pratt & Whitney Canada Corp. Turbine premixing combustor
US20060035183A1 (en) * 2003-02-14 2006-02-16 Richard Carroni Mixer
US20060107647A1 (en) * 2004-04-20 2006-05-25 Labala Gustavo F Turbine, particularly useful for small aircraft
US7065954B2 (en) * 2004-04-20 2006-06-27 Gustavo Francisco Labala Turbine, particularly useful for small aircraft
US7513098B2 (en) 2005-06-29 2009-04-07 Siemens Energy, Inc. Swirler assembly and combinations of same in gas turbine engine combustors
US20070130951A1 (en) * 2005-12-10 2007-06-14 Seoul National University Industry Foundation Combustor
EP2436977A1 (de) * 2010-09-30 2012-04-04 Siemens Aktiengesellschaft Brenner für eine Gasturbine
WO2012041839A1 (en) * 2010-09-30 2012-04-05 Siemens Aktiengesellschaft Burner for a gas turbine
CN103140714A (zh) * 2010-09-30 2013-06-05 西门子公司 用于燃气轮机的燃烧器
US20140102572A1 (en) * 2012-10-17 2014-04-17 Delavan Inc. Radial vane inner air swirlers
US9488108B2 (en) * 2012-10-17 2016-11-08 Delavan Inc. Radial vane inner air swirlers

Also Published As

Publication number Publication date
DE2415036C2 (de) 1982-12-16
DE2415036A1 (de) 1974-11-07
GB1423799A (en) 1976-02-04
JPS5421482B2 (de) 1979-07-31
JPS502222A (de) 1975-01-10
CA982831A (en) 1976-02-03

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