EP0100135B1 - Chambre de combustion - Google Patents

Chambre de combustion Download PDF

Info

Publication number
EP0100135B1
EP0100135B1 EP19830301586 EP83301586A EP0100135B1 EP 0100135 B1 EP0100135 B1 EP 0100135B1 EP 19830301586 EP19830301586 EP 19830301586 EP 83301586 A EP83301586 A EP 83301586A EP 0100135 B1 EP0100135 B1 EP 0100135B1
Authority
EP
European Patent Office
Prior art keywords
air
liner
combustor
combustion
passage
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
Application number
EP19830301586
Other languages
German (de)
English (en)
Other versions
EP0100135A1 (fr
Inventor
Hukam C. Mongia
Edwin B. Coleman
Thomas W. Bruce
Harry A. Elliot
John T. White
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.)
Garrett Corp
Original Assignee
Garrett 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
Priority claimed from US06/400,578 external-priority patent/US4497170A/en
Priority claimed from US06/400,580 external-priority patent/US4532762A/en
Application filed by Garrett Corp filed Critical Garrett Corp
Publication of EP0100135A1 publication Critical patent/EP0100135A1/fr
Application granted granted Critical
Publication of EP0100135B1 publication Critical patent/EP0100135B1/fr
Expired legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow

Definitions

  • the present invention relates generally to combustors for example for use in gas turbine propulsion engines, and one object of the invention is to provide for imparting significantly improved stability and ignition performance to high-temperature-rise combustors.
  • a combustor for example for an aircraft gas turbine engine, which comprises a liner defining a combustion passage, separate first and second air inlet means opening through the liner for the passage of combustion air into the upstream part of the combustion passage, and means for controlling the flow of air through the first air inlet means into the combustion passage
  • the first air inlet means provides separate air flows respectively through air swirler means and liner cooler passage means both situated at an ignition zone in the upstream part of the combustion passage, and in that means are provided for controlling the first air inlet means so as to effect simultaneous variation, in the same incremental or decremental sense, as the case may be, of the air flows respectively through the air swirler means and the liner cooler passage means.
  • the invention thus enables the liner wall temperature in the ignition zone to be controlled in conjunction with fuel mixture control, simply by variation of the air flow admitted through the said first air inlet means.
  • the combustor may include a third air inlet means operating through the liner downstream of the ignition zone for the supply of air to the combustion passage, and means for controlling the first and third air inlet means so as to effect simultaneous control in opposite incremental and decremental senses of the air flows through the first and third air inlet means.
  • a third air inlet means operating through the liner downstream of the ignition zone for the supply of air to the combustion passage, and means for controlling the first and third air inlet means so as to effect simultaneous control in opposite incremental and decremental senses of the air flows through the first and third air inlet means.
  • Each of U.S. patent specifications Nos. 3 919 838, 3 927 520 and 3 930 368, and European patent specification No. 0 026 594 discloses a combustor provided with variable flow control means by which the distribution of combustion air between an upstream reaction zone and a downstream dilution zone of a combustion chamber may be varied, by means of interlinked valves controlling upstream and downstream air inlets into the combustion chamber, the valves operating inversely so that when the upstream valve is opened the downstream valve closes and vice versa.
  • European specification No. 0 026 594 moreover shows such a combustor which has a prechamber into which a fixed amount of air is introduced through a first swirler and mixes with vaporised fuel therein before passing via a flow dam into the upstream end of the combustion chamber for combustion in the reaction zone.
  • a second swirler is provided by means of which a variable proportion of air controlled by one of the interlinked valves is admitted partly into the downstream portion of the prechamber, upstream of the dam, and partly directly into the upstream end of the combustion chamber, bypassing the prechamber and the floor dam, to reduce pressure loss and combustion temperature in the reaction zone.
  • a gas turbine propulsion engine for an aircraft may be provided with such a specially designed combustor, which as explained below is operable to significantly expand the altitude-mach number flight envelope within which the engine may be operated without experiencing the combustor lean instability and relight problems associated with conventional fixed geometry combustors.
  • the combustor has a number of features, which will be briefly discussed in broad terms.
  • that combustor is of an annular, reverse flow configuration, having a hollow, annular combustor liner which is surrounded by an intake plenum that receives high pressure discharge air from the engine's compressor section.
  • the combustor liner has an annular upstream end wall and an annular interior wall spaced therefrom in a downstream direction.
  • annular liner walls Together with the annular liner sidewalls these axially spaced liner walls define within the liner interior an annular air inlet plenum which communicates with the main combustor intake plenum through a circumferentially spaced series of liner slots positioned around the outer liner periphery between the upstream end wall and the interior wall of the liner.
  • the liner inlet plenum and the interior liner wall Projecting axially inwardly through the liner end wall, the liner inlet plenum and the interior liner wall are a circumferentially spaced series of piloted, air blast fuel nozzles for injecting fuel into a dome portion of the liner interior which extends downstream from the annular interior liner wall.
  • Compressor discharge air entering the liner plenum through the slotted inlet openings is forced into the liner dome area through air swirler means carried by the interior wall and circumscribing each of the fuel nozzles. A portion of this entering air is also forced axially through skirted cooling passages extending along the dome portion of the liner interior and communicating with the liner inlet plenum.
  • a circumferentially spaced series of liner wall orifices for admitting primary combustion air into the liner interior.
  • a circumferentially spaced series of aft end openings formed through the radially outer liner sidewall, through which the liner interior and the main combustor intake plenum communicate.
  • means are provided for selectively closing off the liner's slotted inlet plenum openings while permitting air inflow through the aft end openings, or closing off the aft end openings while permitting air inflow through the slotted lines inlet plenum openings.
  • This ability to close off either the aft end openings or the liner plenum inlet openings affords the combustor substantially improved combustion stability and relight capabilities.
  • the aft end openings are unblocked, the liner inlet plenum openings are closed off, and the nozzles are staged to their pilot fuel spray mode.
  • the closure of the liner plenum inlet openings prevents inward airflow through the air swirlers and also prevents the flow of cooling air through the skirted cooling passages.
  • This mode of operating the combustor which is also utilized to effect high altitude relights of combustor, simultaneously maximizes the fuel richness in the liner dome area and minimizes the heat transfer outwardly through the walls thereof.
  • the invention includes a method of operating a combustor in which the supply of air is controlled simultaneously through two inlets.
  • FIG. 1 Schematically illustrated in Fig. 1 are the primary components of a gas turbine propulsion engine 10 which embodies principles of the present invention.
  • ambient air 12 is drawn into a compressor 14 which is spaced apart from and rotationally coupled to a bladed turbine section 16 by an interconnecting shaft 18.
  • Pressurized air 20 discharged from compressor 14 if forced into an annular, reverse flow combustor 22 which circumscribes the turbine section 16 and an adjacent portion of shaft 18.
  • the air 20 is mixed within the combustor with fuel 24, the resulting fuel-air mixture being continuously burned and discharged from the combustor across turbine section 16 in the form of hot, expanded gas 26.
  • This expulsion of the gas 26 simultaneously drives the turbine and compressor, and provides the engine's propulsive thrust.
  • Conventional combustors used in aircraft jet propulsion engines are of fixed geometry construction and are designed to be operated only within a predetermined altitude-mach number flight envelope such as envelope 28 bounded by the solid line 30 in the graph of Fig. 2. If an attempt is made to operate the conventional combustor at higher altitudes or lower mach numbers than those within envelope 28 (i.e., within, for example, the cross-hatched area 32 bounded by line 30 and dashed line 36 in Fig. 2), the lean stability and altitude relight capability of the combustor are adversely affected.
  • the combustor 22 embodying the present invention is of a unique, variable geometry construction which permits the engine 10 to be efficiently and reliably operated within the substantially expanded flight envelope 28, 32 without these lean instability, altitude relight, and ground start problems of fixed geometry combustors.
  • the combustor 22 includes a hollow, annular outer housing 36 having an annular radially outer sidewall 38 and an annular, radially inner sidewall 40 spaced apart from and connected to sidewall 38 by an annular upstream end wall 42. Positioned coaxially within the housing 36 is an upstream end portion of an annular, hollow combustor liner 44 having a reverse flow configuration.
  • Liner 44 has an annular upstream end wall 46 spaced axially inwardly from the housing end wall 42, and annular radially outer and inner sidewalls 48, 50 which extend leftwardly (as viewed in Fig. 3) from liner end wall 46 and then curve radially inwardly through a full 180°.
  • the liner sidewalls 48, 50 define an annular discharge opening 52 through which the hot discharge gas 26 is expelled from the interior or combustion flow passage 54 of liner 44.
  • housing 36 defines an intake plenum 56 which circumscribes the upstream end portion of liner 44 as indicated in Fig. 3.
  • Compressor discharge air 20 is forced into plenum 56 through an annular inlet opening 58 which circumscribes the liner 44 and is positioned at the left end of combustor 22.
  • a portion of this pressurized air is used to cool the liner sidewalls 48, 50 during combustor operation.
  • the sidewalls 48, 50 have, along axially adjacent portions of their lengths, overlapping, radially spaced inner and outer wall segments 48a, 48b and 50a, 50b (only one set of such inner and outer wall segments being representatively illustrated in Fig. 3).
  • air 20 is forced inwardly through openings 49, 51 formed respectively through the wall segments 48b, 50b.
  • the entering air impinges upon the inner wall segments 48a, 50a and enters the combustion flow passage 54, in a downstream direction, through exit slots 48c, 50c formed between the skirted wall segments.
  • annular liner inlet plenum 60 which is positioned axially between the liner end wall 46 and an annular liner interior wall 62 which is axially spaced in a downstream direction from the liner end wall 46.
  • the plenum 60 opens radially outwardly through the outer liner sidewall 48 through a circumferentially spaced series of inlet slots 64 (only one of which is shown in Fig. 3) formed through sidewall 48.
  • Extending downstream from the interior wall 62 is a dome portion 54a of the combustion flow passage 54 which is radially bounded by inner and outer annular cooling skirts 66, 68.
  • Cooling skirts 66, 68 are spaced inwardly from the inner and outer liner sidewalls 50, 48, respectively, and define with the liner sidewalls axially extending cooling passages 70, 72 which open in a downstream direction into the combustion flow passage 54 as indicated in Fig. 3.
  • Cooling passage 70 communicates at its upstream end with the liner inlet plenum 60 through a circumferentially spaced series of air passages 74 formed through the liner interior wall 62, while the annular cooling passage 72 communicates with the plenum 60 through a circumferentially spaced series of air flow passages 76 also extending through the interior wall 62.
  • compressor discharge air 20 is selectively admitted to the liner plenum 60 and is forced axially through the annular flow passages 70, 72 and into the combustion flow passage 54 to thereby cool the inner wall surfaces of the liner dome portion 54a similarly to the cooling of the inner liner wall surfaces achieved by the cooling skirts 48a, 50a.
  • a circumferentially spaced series of fuel nozzles 78 are utilized.
  • the nozzles 78 project axially inwardly through the liner end wall 46, the liner plenum 60, and the liner interior wall 62 into the dome area 54a (see also Fig. 4).
  • Each of these fuel nozzles is of a piloted air blast type, being supplied by a pair of fuel lines 80, 82 extending inwardly through the housing end wall 42.
  • a pressure atomizing fuel outlet (not specifically illustrated) and an air blast fuel spray outlet (also not specifically illustrated).
  • the nozzles may be staged to deliver fuel through either of the atomizing or air blast outlets.
  • each of the nozzles 78 Coannularly circumscribing each of the nozzles 78, and carried by the liner interior wall 62, are a pair of annular air swirlers 84, 86 which provide communication between the liner dome area 54a and the liner plenum 60 radially inwardly of the cooling skirts 66, 68.
  • Primary combustion air is admitted to the flow passage 54 through a circumferentially spaced series of inlet orifices 88 positioned immediately downstream from the dome area 54a.
  • annular plenum 90 which opens outwardly into the housing plenum 56 through a circumferentially spaced series of slots 92 formed through the liner side wall 48, and communicates with the combustion flow passage 54 through a circumferentially spaced series of inlet passages 94 extending inwardly through the sidewall 48.
  • the previously described structure of the combustor 44 uniquely permits its geometry to be effectively varied by selectively blocking or unblocking the inlet slots 64, 92, in a predetermined manner which will now be described, to substantially enhance the lean stability and starting capabilities of the engine 10.
  • a first sealing member in the form of a valve ring 96 is provided.
  • Ring 96 coaxially circumscribes and outwardly overlies an upstream end portion of the combustor liner 44 as best illustrated in Fig. 3.
  • Ring 96 is axially movable relative to the combustor liner between a closed position illustrated in Fig. 3, and an open position illustrated in Fig. 38.
  • a left or forward axial portion 96a of ring 96 is radially outwardly enlarged and has formed therethrough a circumferentially spaced series of inlet slots 98 (Fig. 8).
  • This forward portion 96a of the ring 96 is slidably and sealingly engaged by a piston ring 100 carried by the outer liner wall 48, while the right or rear portion 96b of ring 96 is slidably and sealingly engaged by a piston ring 102 which is carried by the liner end wall 46.
  • a second sealing member 104 is provided for selectively blocking and unblocking the inlet slots 92.
  • Ring 104 coaxially circumscribes and outwardly overlies the liner sidewall 48 for slidable axial movement therealong between a closed position indicated in Fig. 3A and an open position shown in Fig. 3. With the sealing ring 104 in its closed position, the inlet ports 92 are blocked to preclude entry therethrough of compressor discharge air 20, an annular lip 104a on the ring 104 cooperating with an overlying annular lip 106 on the liner side wall 48 to create a labyrinth seal 108 between ring 104 and side wall 48, as shown in Fig. 3A, with ring 104 in its closed position.
  • the rear axial portion 96b thereof blocks off the inlet slots 64 (Fig. 3) to preclude entry of compressor discharge air 20 into the liner inlet plenum 60, the piston rings 100, 102 providing annular air flow seals between the combustor liner and the ring 96 adjacent the opposite ends of the plenum 60.
  • the sealing rings 96, 104 are selectively moved in axially opposite directions (i.e. parallel to the center line or axis 110 of the combustor) between their previously described open and closed positions by a novel actuation system 112.
  • the actuation system includes an actuation or unison ring 114 which is positioned coaxially within the housing plenum 56 immediately adjacent the outer ends of the fuel nozzles 78.
  • the actuation ring 114 is rotatably supported within the plenum 56 by a circumferentially spaced series of bearing support brackets 116 (only one of such brackets being illustrated in Fig. 4) which are positioned between adjacent nozzles 78 and externally secured to the liner end wall 46.
  • Each of these brackets 116 carries a carbon bearing block 118 which is slidably received in a circumferential channel 120 (see Fig. 6) formed in the radially inner surface of the unison ring 114, thereby facilitating rotation of ring 114 within the plenum 56.
  • Control rod 122 extends into a small housing 124, through seal means 126 carried by such housing, which is externally secured to the outer housing sidewall 38 over an opening 128 extending therethrough.
  • Control rod 122- extends lengthwise generally tangentially to the outer surface of housing sidewall 38 and perpendicularly to the combustor axis.
  • the inner end of the control rod 122 is pivotally secured to one end of a connecting rod 130 which extends radially inwardly through the sidewall opening 128 and is secured at its inner end to the unison ring 114. As viewed in Fig.
  • inward axial movement of the control rod 122 which may be achieved by conventional control means (not illustrated) positioned outside the combustor housing, moves the connecting rod 130 to the left within the opening 128 and causes a counter-clockwise rotation of the unison ring 114.
  • an outward axial movement of the control rod causes a clockwise rotation of the unison ring.
  • Such selective rotation of the unison ring 114 is utilized to cause the opposite axial motion of the sealing rings 96,104 by linking means in the form of four circurriferentially spaced sets of actuating rods 132, 134 (only one such rod set being illustrated in Figs. 3 and 8) which extend axially within the housing sidewalls 48, 38 and are connected to the unison ring 114 by means of four circumferentially spaced bell crank members 136.
  • each of the four bell crank members 136 has a base leg portion 138 which is pivoted at its outer end to the unison ring 114 (as at 140) and extends from its pivot point, in a generally axial direction toward the housing end wall 42, to a radially outwardly directed trunk portion 142 which is pivoted in a support bracket 144 as indicated in phantom in Fig. 8.
  • Each of the four support brackets 144 is secured to the liner end wall 46 between an adjacent pair of nozzles 78 as can be best seen in Fig. 4.
  • each of these support brackets 144 also carries a carbon bearing block 118 (see Fig. 3) which slidably engages the inner surface of the unision ring 114.
  • each arm 146 Extending transversely in opposite directions from the bell crank member's trunk portion 142 are a pair of control arms 146, 148 (Fig. 8).
  • the outer end of each arm 146 is pivotally connected to one end of an actuating rod 132 which is secured at its opposite end to the sealing ring 104.
  • the outer end of each arm 148 is pivotally connected to one end of an actuating rod 134, the other end of such actuating rod 134 being secured to the sealing ring 96.
  • each of them is slidably extended through a journal portion 144a of its associated support bracket 144 (Fig. 5) and an additional journal support 150 (Fig. 3) carried by the outer housing sidewall 38.
  • Such journalling also rotationally stabilizes the sealing ring 104, thereby assuring a smooth sliding motion thereof along the liner sidewall 48.
  • a similar rotational stability is also provided to the sealing ring 96 by means of a channel 152 (Fig. 5) which is formed in a guide member 154 secured to the sealing ring 96, the channel 152 slidably receiving a downturned lip portion 156 of support bracket 144 (see also Fig. 4).
  • the actuation system 112 is utilized to move the sealing ring 96 to its open position (Fig. 3B) and to simultaneously move the sealing ring 104 to its closed position (Fig. 3A).
  • compressor discharge air 20 in the housing plenum 56 is forced inwardly through the sealing ring inlet slots 98 into the liner plenum 60.
  • From the plenum 60 entering air 20 is forced outwardly through the dome wall cooling slots 70, 72 and is also forced into the combustor dome portion 54a through the annular air swirlers 84, 86 in a swirling flow pattern.
  • the entering swirl air is mixed with the fuel and fuel-air mixtures discharged from the nozzles 78, further mixed with the primary combustion air entering through the primary orifices 88, and continously burned.
  • the fuel richness within the combustor dome area 54a may be selectively varied both by variably staging the fuel nozzles 78 and by moving the sealing ring 96 toward its closed position, thereby blocking off a portion of the sealing ring inlet slots 98.
  • Such movement of the sealing ring 96 toward its closed position simultaneously reduces air flow through the wall cooling slots 70, 72 and through the swirler plates 84, 86. This, in turn, reduces the dome wall cooling, thereby elevating the combustion temperature in the dome area, and reduces the total amount of swirler air entering the dome area.
  • the overall combustion stability of the combustor 22 is substantially improved compared to conventional fixed geometry combustors, thus permitting reliable and efficient normal operation of the engine 10 within the expanded flight envelope portion 32 of Fig. 2.
  • the altitude restart capabilities of the combustor 22 are further enhanced when the sealing ring 104 is brought to its fully opened positioned by the actuation system 112.
  • excess compressor discharge air is intentionally bypassed around the combustor and bled off to atmosphere.
  • the present invention such compressor discharge air is uniquely utilised to assist in the altitude restart procedure. More specifically, with the sealing ring 104 in its fully opened position, this previously wasted excess compressor discharge air is forced imwardly through the inlet slots 92, the plenum 90 and the inlet passages 94 into the combustion flow passage 54. The entering compressor discharge air is then forced outwardly through the combustor outlet opening 52 and across the bladed turbine section 16 to greatly assist in the "windmill" restarting of the engine 10.
  • the sealing ring 104 is in its closed position so that the air for combustion is not extracted through the passages 94.
  • maximisation of the fuel richness and wall temperatures within the dome area 54a not only improves the altitude relight and lean stability characteristics of the engine 10 but also substantially improves its ground start capabilities-especially in low ambient temperature conditions.
  • the present invention provides improved combustor apparatus, and associated operating methods, which eliminate or substantially reduce the stability and relight problems commonly associated with conventional fixed geometry combustors.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Claims (9)

1. Chambre de combustion, par exemple pour un turbo-moteur d'aviation, comportant une enveloppe (44) qui definit un passage de combustion (54), cette enveloppe étant pourvue d'orifices d'admission d'air d'un premier genre (64) et d'un second genre (88), menagés séparément à travers l'enveloppe, pour introduire l'air de combustion dans la partie amont du passage de combustion, et de moyens (96, 98, 134) pour régler le débit de l'air qui pénètre dans le passage de combustion, à travers les orifices d'admission d'air du premier genre (64); caractérisée en ce que les orifices d'admission d'air du premier genre (64) assurent l'écoulement de courants d'air séparés, respectivement à travers des passages de turbulence (84, 86) et à travers des passages (70, 72) prévus pour le refroidissement de l'enveloppe; ces deux groupes de passages étant situés à l'endroit d'une zone d'allumage (54a), dans la partie amont du passage de combustion; des moyens de réglage (96, 98, 134) étant prévus, pour agir sur les orifices d'admission d'air du premier genre (64) de manière à faire varier simultanément et dans le même sens, en augmentation ou en diminution suivant le cas, le débit d'air à travers les passages de turbulence et à travers les passages de refroidissement de l'enveloppe, respectivement.
2. Chambre de combustion selon la revendication 1, caractérisée en ce qu'elle comporte des orifices d'admission d'air d'un troisième genre (92, 94), qui débouchent à travers l'enveloppe en aval de la zone d'allumage, pour assurer un débit d'air dans le passage de combustion, et des moyens (112) pour agir sur les orifices d'admission d'air du premier genre (64) et du troisième genre (92, 94) de manière à faire varier simultanément et en sens opposés, en augmentation ou en diminution, le débit d'air à travers les orifics d'admission d'air du premier genre et du troisième genre.
3. Chambre de combustion selon l'une des revendications 1 ou 2, comportant des moyens d'actionnement (122, 130, etc.) pour commander les organes de réglage (96, 98,134, ou 96, 98,134, 112) des orifices d'admission d'air du premier genre (64), ou le cas échéant des orifices d'admission d'air du premier genre (64) et du troisième genre (92, 94); ces moyens d'actionnement comportant une couronne de con- j□gaison (114) montée de manière à tourner suivant un axe de la chambre de combustion, et des moyens de liaison (figure 8), pour déplacer axialement les organes de réglage asservis à la rotation de la couronne de conjugaison (114).
4. Chambre de combustion selon l'une des revendications 2 ou 3, caractérisée en ce que les moyens de liaison comportent une série de systèmes de leviers coudés de renvoi (142), reliés chacun par une articulation à la couronne de conjugaison (114), des moyens pour monter les systèmes de leviers de renvoi sur l'enveloppe de manière pivotante suivant des axes sensiblement perpendiculaires à l'axe de la chambre de combustion; un premier groue de tiges d'actionnement (134) réparties autour de l'enveloppe, et ayant leurs extrémités opposées reliées aux organes de réglage (96) des orifices d'admission d'air du premier genre (64) et à l'un des systèmes de leviers coudés de renvoi; et un second groupe de tiges d'actionnement (132) réparties autour de l'enveloppe, et ayant leurs extrémités opposées reliées aux organes de réglage (104) des orifices d'admission d'air du troisième genre (92, 94) et a l'un des systèmes de leviers coudés de renvoi; et un organe de manoeuvre mobile (122, 130) qui pénètre dans le corps de la chambre de combustion, pour commander sélectivement la rotation de la couronne de conjugaison (114).
5. Chambre de combustion selon l'une des revendiGations 1 à 4, caractérisée en ce qu'elle comporte une cavité d'admission interne (60), disposée dans l'enveloppe de la chambre de combustion à l'endroit de l'extrémité amont du passage de combustion, pour recevoir le débit d'air réglé par les organes de réglage (96) des orifices d'admission d'air du premier genre (64), et pour envoyer ce débit d'air dans l'extrémité amont du passage de combustion (54), à travers les passages de turbulence (84, 86) et les passages de refroidissement (70, 72) de l'enveloppe.
6. Chambre de combustion selon l'une des revendications 1 à 5, caractérisée en ce que l'enveloppe (44) comporte un fond annulaire (46), disposé à l'intérieur d'un corps creux (36) pour définir dans celui-ci une enceinte d'admission d'air (56).
7. Chambre de combustion selon la revendication 6, caractérisée en ce qu'elle comporte un groupe de gicleurs de carburant (78) montés dans le fond annulaire (46) de l'enveloppe et un groupe d'orifices d'entrée d'air primaire (88) ménagés dans la paroi latérale de l'enveloppe.
8. Procédé de mise en oeuvre d'une chambre de combustion, caractérisé en ce qu'il comporte les phases opératoires suivantes: (a) on introduit un courant d'air sous pression provenant d'une source d'air sous pression (56) dans une enveloppe (48) de la chambre de combustion, par une ouverture d'admission d'air (64) ménagée dans cette enveloppe; (b) on utilise une première partie du courant d'air sous pression qui pénètre dans l'enveloppe, pour refroidir une partie de la surface interne de cette enveloppe; (c) on utilise une seconde partie du courant d'air sous pression qui pénètre dans l'enveloppe, comme air de combustion dans cette enveloppe; (d) on mélange cette seconde partie de l'air avec un carburant (24), et on brûle ce mélange de carburant et d'air dans l'enveloppe; et (e) on régie en même temps le refroidissement de ladite partie de la surface interne de l'enveloppe, et la richesse du mélange d'air et de carburant, en faisant varier le débit de l'air qui passe à travers l'ouverture d'admission d'air (64).
9. Procédé de mise en oeuvre d'un moteur d'aviation à turbine à gaz, comportant un compresseur, une turbine et une chambre de combustion; cette chambre de combustion comportant un corps creux dont la paroi extérieure (38) entoure au moins en partie une enveloppe (48) de la chambre de combustion, en definissant dans ce corps creux une enceinte d'admission (56), pour recevoir l'air comprimé refoulé par le compresseur; cette enveloppe de la chambre de combustion définissant un passage interne (54) pour l'écoulement des gaz de combustion, et comportant un premier genre (64) et un second genre (92) d'orifices séparés d'admission d'air, pour l'introduction de l'air sous pression de l'enceinte d'admission dans le passage d'écoulement des gaz de combustion; le procédé étant caractérisé en ce qu'il comporte les phases opératoires suivantes: (a) on fait passer vers l'intérieur un courant d'air sous pression de t'enceinte admission (56) du corps creux, à travers les orifices d'admission d'air du premier genre (64), tout en obturant les orifices d'admission d'air du second genre (92), pour un régime de fonctionnement normal du moteur; (b) on utilise le courant d'air qui passe à travers les orifices d'admission du premier genre, pour assurer en même temps une partie de la surface interne de l'enveloppe de la chambre de combustion, et pour fournir l'air de combustion dans le passage d'écoulement des gaz de combustion; (c) on fait passer vers l'intérieur un courant d'air sous pression de l'enceinte d'admission du corps creux, à travers les orifices d'admission du second genre (92), tout en obturant les orifices d'admission du premier genre (64) pour le démarrage du moteur; et (e) on utilise le courant d'air qui passe à travers les orifices d'admission d'air du second genre (92), pour le faire participer à l'entraînement de la turbine pendant le démarrage du moteur.
EP19830301586 1982-07-22 1983-03-22 Chambre de combustion Expired EP0100135B1 (fr)

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
US400578 1982-07-22
US06/400,578 US4497170A (en) 1982-07-22 1982-07-22 Actuation system for a variable geometry combustor
US06/400,580 US4532762A (en) 1982-07-22 1982-07-22 Gas turbine engine variable geometry combustor apparatus
US400580 1982-07-22

Publications (2)

Publication Number Publication Date
EP0100135A1 EP0100135A1 (fr) 1984-02-08
EP0100135B1 true EP0100135B1 (fr) 1986-06-11

Family

ID=27017100

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19830301586 Expired EP0100135B1 (fr) 1982-07-22 1983-03-22 Chambre de combustion

Country Status (2)

Country Link
EP (1) EP0100135B1 (fr)
DE (1) DE3364029D1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2112181C1 (ru) * 1994-06-10 1998-05-27 Ставропольское высшее авиационное инженерное училище противовоздушной обороны им.Маршала авиации Судца В.А. Камера сгорания газотурбинного двигателя

Families Citing this family (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3942451A1 (de) * 1989-12-22 1991-06-27 Daimler Benz Ag Vorrichtung zur einstellung der sekundaerluftmenge an einer gasturbinenbrennkammer
FR2690977B1 (fr) * 1992-05-06 1995-09-01 Snecma Chambre de combustion comportant des passages reglables d'admission de comburant primaire.
FR2704628B1 (fr) * 1993-04-29 1995-06-09 Snecma Chambre de combustion comportant un système d'injection de comburant à géométrie variable.
US6220034B1 (en) 1993-07-07 2001-04-24 R. Jan Mowill Convectively cooled, single stage, fully premixed controllable fuel/air combustor
US5572862A (en) * 1993-07-07 1996-11-12 Mowill Rolf Jan Convectively cooled, single stage, fully premixed fuel/air combustor for gas turbine engine modules
US5638674A (en) * 1993-07-07 1997-06-17 Mowill; R. Jan Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission
RU2163991C2 (ru) * 1995-12-19 2001-03-10 Акционерное общество открытого типа "Самарский научно-технический комплекс "Двигатели НК" Камера сгорания газотурбинного двигателя с регулируемым распределением воздуха
RU2158881C2 (ru) * 1996-01-10 2000-11-10 Ставропольское высшее авиационное инженерное училище ПВО им.маршала авиации Судца В.А. Фронтовое устройство камеры сгорания для высокотемпературного гтд
RU2159898C2 (ru) * 1996-01-10 2000-11-27 Ставропольское высшее авиационное инженерное училище ПВО им.маршала авиации В.А.Судца Диффузор основной камеры сгорания авиационного газотурбинного двигателя
US5924276A (en) * 1996-07-17 1999-07-20 Mowill; R. Jan Premixer with dilution air bypass valve assembly
FR2780488B1 (fr) * 1998-06-24 2000-10-13 Pillard Chauffage Amelioration aux appareils a combustion comportant plusieurs conduits de transport de comburant
RU2286513C1 (ru) * 2005-05-05 2006-10-27 Сергей Александрович Маяцкий Устройство для сжигания топлива в газотурбинном двигателе
RU2311589C1 (ru) * 2006-05-04 2007-11-27 Сергей Александрович Маяцкий Камера сгорания газотурбинного двигателя
RU2325588C2 (ru) * 2006-07-13 2008-05-27 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения имени П.И. Баранова" Устройство для регулирования низкоэмиссионной камеры сгорания газовой турбины
RU2376528C1 (ru) * 2008-03-31 2009-12-20 Денис Сергеевич Легконогих Способ повышения живучести камеры сгорания газотурбинного двигателя и камера сгорания
RU2378576C1 (ru) * 2008-04-02 2010-01-10 Закрытое акционерное общество "Энергомаш (Холдинг)" Горелочное устройство камеры сгорания газотурбинной установки
RU2400673C1 (ru) * 2009-01-11 2010-09-27 Открытое акционерное общество "Климов" Камера сгорания с оптимальным режимом работы
RU2505749C1 (ru) * 2012-07-27 2014-01-27 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения им. П.И. Баранова" Камера сгорания газотурбинного двигателя и способ ее работы
RU2513527C1 (ru) * 2012-12-20 2014-04-20 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения им. П.И. Баранова" Камера сгорания газотурбинного двигателя и способ ее работы
RU2531477C1 (ru) * 2013-08-30 2014-10-20 Федеральное государственное казенное военное образовательное учреждение высшего профессионального образования "Военный учебно-научный центр Военно-воздушных сил "Военно-воздушная академия имени профессора Н.Е. Жуковского и Ю.А. Гагарина" (г. Воронеж) Министерства обороны Российской Федерации Устройство для сжигания топлива в газотурбинном двигателе
RU2625076C1 (ru) * 2016-02-08 2017-07-11 Николай Борисович Болотин Камера сгорания газотурбинного двигателя и средство активации воздуха
CN115949509B (zh) * 2023-01-09 2025-09-05 中国航发湖南动力机械研究所 一种点火起动装置及燃气轮机

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3927520A (en) * 1974-02-04 1975-12-23 Gen Motors Corp Combustion apparatus with combustion and dilution air modulating means
US3919838A (en) * 1974-11-04 1975-11-18 Gen Motors Corp Combustion control
US3930368A (en) * 1974-12-12 1976-01-06 General Motors Corporation Combustion liner air valve
US4263780A (en) * 1979-09-28 1981-04-28 General Motors Corporation Lean prechamber outflow combustor with sets of primary air entrances

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2112181C1 (ru) * 1994-06-10 1998-05-27 Ставропольское высшее авиационное инженерное училище противовоздушной обороны им.Маршала авиации Судца В.А. Камера сгорания газотурбинного двигателя

Also Published As

Publication number Publication date
DE3364029D1 (en) 1986-07-17
EP0100135A1 (fr) 1984-02-08

Similar Documents

Publication Publication Date Title
EP0100135B1 (fr) Chambre de combustion
US4497170A (en) Actuation system for a variable geometry combustor
US4532762A (en) Gas turbine engine variable geometry combustor apparatus
US4567724A (en) Variable geometry combustor apparatus and associated methods
US8011188B2 (en) Augmentor with trapped vortex cavity pilot
US5694767A (en) Variable slot bypass injector system
US4112676A (en) Hybrid combustor with staged injection of pre-mixed fuel
US5664412A (en) Variable geometry air-fuel injector
US4899539A (en) Flow mixer and flame stabilizer for a turbofan engine
US4043121A (en) Two-spool variable cycle engine
EP0644994B1 (fr) Appareil a chambre de combustion et procede pour effectuer la combustion
US6199367B1 (en) Air modulated carburetor with axially moveable fuel injector tip and swirler assembly responsive to fuel pressure
US6983601B2 (en) Method and apparatus for gas turbine engines
US4594848A (en) Gas turbine combustor operating method
US4817378A (en) Gas turbine engine with augmentor and variable area bypass injector
CA2567533C (fr) Dispositif d'augmentation de puissance pour cavite de turbulence enfermee et ravitaillee en carburant de l'exterieur
US6981361B2 (en) Methods for operating gas turbine engines
US5309710A (en) Gas turbine combustor having poppet valves for air distribution control
US4586328A (en) Combustion apparatus including an air-fuel premixing chamber
EP2400221B1 (fr) Purge d'éjecteur de conduit d'écoulement d'échappement adjacent de cavité
US4502636A (en) Variable geometry ejector nozzle for turbomachines
JP3030041B2 (ja) 燃焼器
US4170109A (en) Thrust augmentor having swirled flows for combustion stabilization
GB2503184A (en) Combustor
JPH06281145A (ja) ガスタービン燃焼器およびその燃焼用空気量制御方法

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Designated state(s): DE FR GB IT SE

17P Request for examination filed

Effective date: 19840727

ITF It: translation for a ep patent filed
GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE FR GB IT SE

REF Corresponds to:

Ref document number: 3364029

Country of ref document: DE

Date of ref document: 19860717

ET Fr: translation filed
PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
ITTA It: last paid annual fee
PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 19911223

Year of fee payment: 10

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 19920312

Year of fee payment: 10

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: SE

Payment date: 19920319

Year of fee payment: 10

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 19920430

Year of fee payment: 10

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Effective date: 19930322

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Effective date: 19930323

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 19930322

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Effective date: 19931130

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Effective date: 19931201

REG Reference to a national code

Ref country code: FR

Ref legal event code: ST

EUG Se: european patent has lapsed

Ref document number: 83301586.0

Effective date: 19931008