EP2003399A2 - Turbomachine combustion chamber with helical air circulation - Google Patents

Turbomachine combustion chamber with helical air circulation Download PDF

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Publication number
EP2003399A2
EP2003399A2 EP08158059A EP08158059A EP2003399A2 EP 2003399 A2 EP2003399 A2 EP 2003399A2 EP 08158059 A EP08158059 A EP 08158059A EP 08158059 A EP08158059 A EP 08158059A EP 2003399 A2 EP2003399 A2 EP 2003399A2
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EP
European Patent Office
Prior art keywords
combustion chamber
pilot
wall
air
turbomachine
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.)
Granted
Application number
EP08158059A
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German (de)
French (fr)
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EP2003399A3 (en
EP2003399B1 (en
Inventor
Laurent Bernard Cameriano
Michel Pierre Cazalens
Sylvain Duval
Romain Nicolas Lunel
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Safran Aircraft Engines SAS
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SNECMA SAS
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Publication of EP2003399A3 publication Critical patent/EP2003399A3/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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing

Definitions

  • the present invention relates to the general field of combustion chambers of an aeronautical or terrestrial turbomachine.
  • An aeronautical or terrestrial turbomachine is typically formed of an assembly comprising in particular an annular compression section intended to compress air passing through the turbomachine, an annular combustion section disposed at the outlet of the compression section and in which the air coming from the compression section is mixed with fuel for burning, and an annular turbine section disposed at the outlet of the combustion section and a rotor is rotated by gases from the combustion section.
  • the compression section is in the form of a plurality of stages of movable wheels each carrying blades which are arranged in an annular channel through which the air of the turbomachine and whose section decreases from upstream to downstream.
  • the combustion section includes a combustion chamber in the form of an annular channel in which the compressed air is mixed with fuel for burning.
  • the turbine section it is formed by a plurality of stages of moving wheels each carrying blades which are arranged in an annular channel through which the combustion gases pass.
  • the circulation of air through this assembly is generally carried out as follows: the compressed air from the last stage of the compression section has a natural rotational movement with an inclination of the order of 35 ° to 45 ° ° with respect to the longitudinal axis of the turbomachine, tilt which varies according to the speed of the turbomachine (speed of rotation).
  • this compressed air is straightened in the longitudinal axis of the turbomachine (that is to say that the inclination of the air with respect to the longitudinal axis of the turbomachine is brought back at 0 °) via an air rectifier.
  • the air in the combustion chamber is then mixed with fuel so as to ensure satisfactory combustion and the gases from this combustion continue a course generally along the longitudinal axis of the turbomachine to reach the turbine section.
  • the combustion gases are reoriented by a distributor to present a gyratory movement with an inclination greater than 70 ° with respect to the longitudinal axis of the turbine engine.
  • Such inclination is essential to produce the angle of attack required for the mechanical force driving in rotation of the moving wheel of the first stage of the turbine section.
  • Such angular distribution of the air passing through the turbomachine has many disadvantages. Indeed, the air that naturally leaves the last stage of the compression section with an angle between 35 ° and 45 ° is successively rectified (angle reduced to 0 °) at its entry into the combustion section and then reoriented with an angle greater than 70 ° at its entry into the turbine section. These successive angular modifications of the distribution of air through the turbomachine require intense aerodynamic efforts produced by the rectifier of the compression section and the distributor of the turbine section, aerodynamic forces which are particularly detrimental to the overall efficiency of the turbomachine. the turbomachine.
  • the present invention aims to overcome the aforementioned drawbacks by providing a turbomachine combustion chamber that can be powered by an air that has a rotational movement with respect to the longitudinal axis of the turbomachine.
  • the combustion chamber according to the invention can be supplied with air having a rotational movement about the longitudinal axis of the turbomachine.
  • the natural inclination of the air at the outlet of the compression section of the turbomachine can therefore be maintained through the combustion chamber.
  • the aerodynamic force required for rotating the first stage of the turbine section of the turbomachine is considerably reduced. This sharp decrease in aerodynamic forces generates a gain in efficiency of the turbomachine.
  • the rectifier of the compression section and the distributor of the turbine section can be simplified or even eliminated, which represents a saving in weight and a reduction in production costs.
  • pilot cavities which are carburized only for the idling speeds of the turbomachine, provides a stabilization of the combustion flame for all operating conditions of the turbomachine.
  • each pilot cavity is closed at its upstream end and open at its downstream end.
  • each pilot cavity is delimited circumferentially by two substantially radial partitions, one of these partitions comprising a plurality of air injection orifices opening towards the outside of the combustion chamber and opening in said pilot cavity.
  • the other partition of each pilot cavity has, in cross section, a substantially curvilinear section.
  • the full-throttle injectors are offset axially downstream relative to the pilot injectors. Indeed, the flame from the pilot injectors needs a residence time in the combustion chamber which is higher than the flame from the injectors full throttle.
  • the combustion chamber may be devoid of wall connecting transversely upstream longitudinal ends of the inner and outer walls.
  • the absence of such a wall makes it possible to preserve as much as possible the rotation of the air coming from the compression section of the turbomachine.
  • the fuel injection systems are devoid of associated air systems.
  • the combustion chamber may further comprise an inner annular fairing which is mounted on the inner wall in the upstream extension thereof and an outer annular fairing which is mounted on the outer wall in the upstream extension thereof.
  • the invention also relates to a turbomachine comprising a combustion chamber as defined above.
  • the turbomachine partially shown on the figure 1 has a longitudinal axis XX. According to this axis, it comprises in particular: an annular compression section 100, an annular combustion section 200 disposed at the outlet of the section of compression 100 according to the direction of flow of the air passing through the turbomachine, and an annular turbine section 300 disposed at the outlet of the combustion section 200.
  • the air injected into the turbomachine therefore passes successively through the compression section 100, then the combustion section 200 and finally the turbine section 300.
  • the compression section 100 is in the form of a plurality of stages of movable wheels 102 each carrying blades 104 (only the last stage of the compression section is shown in FIG. figure 1 ).
  • the blades 104 of these stages are disposed in an annular channel 106 through which air flows through the turbomachine and whose section decreases from upstream to downstream. Thus, as the air injected into the turbomachine passes through the compression section, it is more and more compressed.
  • the combustion section 200 is also in the form of an annular channel in which the compressed air from the compression section 100 is mixed with fuel for burning there.
  • the combustion section comprises a combustion chamber 202 inside which is burned the air / fuel mixture (this chamber is detailed later).
  • the combustion section 200 also comprises a turbomachine casing formed of an outer annular casing 204 centered on the longitudinal axis XX of the turbomachine and an inner annular casing 206 which is fixed coaxially inside the casing. outer envelope. An annular space 208 formed between these two envelopes 204, 206 receives compressed air from the compression section 100 of the turbomachine.
  • the turbomachine section 300 of the turbomachine is formed by a plurality of stages of movable wheels 302 each carrying blades 304 (only the first stage of the turbine section is shown in FIG. figure 1 ).
  • the blades 304 of these stages are arranged in an annular channel 306 traversed by the gases coming from the combustion section 200.
  • the gases coming from the combustion section must have an inclination relative to the longitudinal axis XX of the turbomachine which is sufficient to rotate the different stages of the turbine section. turbine.
  • a distributor 308 is mounted directly downstream of the combustion chamber 202 and upstream of the first stage 302 of the turbine section 300.
  • This distributor 308 consists of a plurality of fixed radial vanes 310 of which inclination with respect to the longitudinal axis XX of the turbomachine makes it possible to give the gases coming from the combustion section 200 the inclination necessary for driving in rotation the different stages of the turbine section.
  • the distribution of the air successively passing through the compression section 100, the combustion section 200 and the turbine section 300 takes place as follows.
  • the compressed air from the last stage 102 of the compression section 100 naturally has a gyratory movement with an inclination of the order of 35 ° to 45 ° relative to the longitudinal axis X-X of the turbomachine.
  • this inclination angle is reduced to 0 °.
  • the gases resulting from the combustion are redirected by the blades 310 of the distributor 308 of the latter to give them a gyratory movement with an inclination with respect to the longitudinal axis XX which is greater than 70 °.
  • a new architecture of the combustion chamber 202 which can be powered by an air having a rotational movement about the longitudinal axis X-X of the turbomachine.
  • an architecture it is possible to maintain the natural inclination of the compressed air from the last stage of the compression section without having to straighten it in the longitudinal axis X-X.
  • the stationary blades 310 of the distributor 308 of the turbine section 300 it is no longer necessary for the stationary blades 310 of the distributor 308 of the turbine section 300 to have such a large inclination to produce the angle of attack required for the mechanical driving force in rotation of the moving wheel. 302 of the first stage of the turbine section.
  • the combustion chamber 202 comprises an inner annular wall 212 centered on the longitudinal axis XX of the turbomachine, and an outer annular wall 214 also centered on the longitudinal axis XX and surrounding the inner wall of the engine. to define with it an annular space 216 forming a combustion chamber.
  • the combustion chamber 202 further comprises at least one air inlet opening 218 which opens into the combustion chamber 216 at the upstream end thereof and in a substantially longitudinal direction.
  • the section of this air intake opening is adapted to ensure the operation of the combustion chamber.
  • this air inlet opening 218 is formed between the upstream ends of the inner and outer walls 212 and 214 of the combustion chamber.
  • the combustion chamber 202 also comprises a plurality of fuel injection systems 220 distributed on the outer wall 214 around the longitudinal axis XX of the turbomachine and opening into the combustion chamber 216 in a substantially radial direction .
  • the fuel injection systems 220 comprise pilot injectors 220a circumferentially alternating with full-throttle injectors 220b, the full-throttle injectors preferably being offset axially downstream relative to the pilot injectors.
  • the pilot injectors 220a provide ignition and idle phases of the turbomachine and the 220b full-throttle injectors are involved in the take-off, climb and cruise phases.
  • the pilot injectors are fueled continuously while the full-throttle injectors are only fed beyond a certain determined speed.
  • the fuel injection systems 220 are devoid of associated air systems such as air swirlers which make it possible, in a manner known per se, to generate a rotary air flow. inside the combustion chamber in order to stabilize the combustion flame.
  • pilot and full throttle injectors of the combustion chamber are of very simple design and very reliable operation since they are reduced to their simplest function, namely to inject fuel.
  • the pilot injectors 220a are of the same type as the full throttle injectors 220b.
  • the outer wall 214 of the combustion chamber comprises a plurality of pilot cavities 222 which are regularly distributed around the longitudinal axis X-X.
  • each pilot cavity 222 extends, firstly longitudinally between the two longitudinal ends (upstream and downstream) of the outer wall 214, and secondly radially outwardly thereof.
  • the outer wall 214 is profiled with a plurality of cavities 222 protruding outwardly from the wall.
  • pilot cavities 222 are each delimited circumferentially by two partitions 224 which each project radially outwardly with respect to the outer wall 214. As shown in FIGS. figures 2 and 5 , one of these partitions has a plurality of air injection orifices 226 which make it possible to inject air outside the combustion chamber into the pilot cavity in a circumferential direction.
  • the circumferential injection of air is performed in the same direction of rotation (that of the needles of a watch for the exemplary embodiment of figures 2 and 3 ) for all the pilot cavities 222 of the combustion chamber. Moreover, the direction of rotation for the circumferential injection of air into these pilot cavities is that of the compressed air coming from the compression section of the turbomachine.
  • pilot cavities 222 are supplied with fuel via pilot injectors 220a, each of which opens radially into one of these cavities. As for the full-throttle injectors 220b, they each open radially into the combustion chamber between two adjacent pilot cavities.
  • Each pilot cavity 222 is preferably closed at its upstream end by a radial partition 228 and open at its downstream end (see especially the figures 2 and 5 ). Thus, the air entering the combustion chamber 216 through its air inlet opening 218 does not disturb the flow of air introduced into the pilot cavities 222 by the air injection orifices 226.
  • the operation of the combustion chamber is as follows: the compressed air coming from the compression section 100 and which is rotated about the longitudinal axis XX enters the combustion section 200. This air is divided into two flows: “internal” flow and "external” flow.
  • the external flow bypasses the combustion chamber 202 and feeds the pilot cavities 222 after cooling the outer wall 214 of the combustion chamber and the outer casing 204 of the combustion section. This external air is injected into these pilot cavities via the air injection orifices 226 in the direction of rotation of the air at its entry into the combustion section. In these pilot cavities, the air is mixed and burnt with the fuel injected by the pilot injectors 220a.
  • the internal flow that represents the main flow it enters the combustion chamber 216 through the air inlet opening 218 to be mixed and burned fuel injected by the full-throttle injectors 220b. Stabilization of the combustion flame is obtained thanks to the "carburation" of the pilot cavities.
  • each pilot cavity 222 which is not provided with air injection orifices has, in cross-section, a substantially curvilinear section (unlike the other wall which is substantially flat).
  • the curvature of these walls makes it possible to accompany the rotational movement of the air injected into the pilot cavities by the air injection orifices 226.
  • the two longitudinal partitions 224 circumferentially delimiting each pilot cavity 222 are substantially flat and each extend in a radial direction.
  • the number and the geometric dimensions of the pilot cavities 222 of the combustion chamber may vary according to the needs. The same is true of the number, the dimensions and the positioning of the air injection orifices 226 in these cavities.
  • the combustion chamber 202 may also comprise an internal annular fairing 230 which is mounted on the inner wall 212 in the upstream extension thereof and an outer annular fairing 232 which is mounted on the outer wall 214 in the upstream extension thereof.
  • the presence of these shrouds 230, 232 makes it possible to regulate the flow rate of air entering the combustion chamber 202 and that bypassing it.
  • the outer wall 214 of the combustion chamber may comprise at its downstream end an annular flange 234 extending radially outwardly of the wall, this flange being provided with a plurality of holes 236 regularly distributed around the longitudinal axis XX and intended to supply cooling air to the turbine section 300.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Fuel-Injection Apparatus (AREA)
  • Supercharger (AREA)

Abstract

The chamber (202) has an air inlet opening emerging from a combustion hearth, and an outer wall (214) with pilot cavities (222), where each cavity is extended between wall's longitudinal upstream and downstream ends and radially towards the wall's exterior. The cavities are supplied by air exterior to the chamber along circumferential direction. A fuel injection system (220) has pilot injectors (220a) circumferentially alternating with open gas injectors (220b), where each injector (220a) is radially emerged in one cavity. Each injector (220b) is radially emerged between adjacent cavities.

Description

La présente invention se rapporte au domaine général des chambres de combustion d'une turbomachine aéronautique ou terrestre.The present invention relates to the general field of combustion chambers of an aeronautical or terrestrial turbomachine.

Une turbomachine aéronautique ou terrestre est typiquement formée d'un ensemble comportant notamment une section annulaire de compression destinée à comprimer de l'air traversant la turbomachine, une section annulaire de combustion disposée en sortie de la section de compression et dans laquelle l'air issu de la section de compression est mélangé à du carburant pour y être brûlé, et une section annulaire de turbine disposée en sortie de la section de combustion et dont un rotor est entraîné en rotation par des gaz issus de la section de combustion.An aeronautical or terrestrial turbomachine is typically formed of an assembly comprising in particular an annular compression section intended to compress air passing through the turbomachine, an annular combustion section disposed at the outlet of the compression section and in which the air coming from the compression section is mixed with fuel for burning, and an annular turbine section disposed at the outlet of the combustion section and a rotor is rotated by gases from the combustion section.

La section de compression se présente sous la forme d'une pluralité d'étages de roues mobiles portant chacune des aubes qui sont disposées dans un canal annulaire traversé par l'air de la turbomachine et dont la section diminue d'amont en aval. La section de combustion comprend une chambre de combustion se présentant sous la forme d'un canal annulaire dans lequel l'air comprimé est mélangé à du carburant pour y être brûlé. Quant à la section de turbine, elle est formée par une pluralité d'étages de roues mobiles portant chacune des aubes qui sont disposées dans un canal annulaire traversé par les gaz de combustion.The compression section is in the form of a plurality of stages of movable wheels each carrying blades which are arranged in an annular channel through which the air of the turbomachine and whose section decreases from upstream to downstream. The combustion section includes a combustion chamber in the form of an annular channel in which the compressed air is mixed with fuel for burning. As for the turbine section, it is formed by a plurality of stages of moving wheels each carrying blades which are arranged in an annular channel through which the combustion gases pass.

La circulation de l'air au travers de cet ensemble s'effectue généralement de la manière suivante : l'air comprimé issu du dernier étage de la section de compression possède un mouvement giratoire naturel avec une inclinaison de l'ordre de 35° à 45° par rapport à l'axe longitudinal de la turbomachine, inclinaison qui varie en fonction du régime de la turbomachine (vitesse de rotation). A son entrée dans la section de combustion, cet air comprimé est redressé dans l'axe longitudinal de la turbomachine (c'est-à-dire que l'inclinaison de l'air par rapport à l'axe longitudinal de la turbomachine est ramenée à 0°) par l'intermédiaire d'un redresseur d'air. L'air dans la chambre de combustion est alors mélangé à du carburant de manière à assurer une combustion satisfaisante et les gaz issus de cette combustion poursuivent un parcours globalement selon l'axe longitudinal de la turbomachine pour parvenir à la section de turbine. Au niveau de cette dernière, les gaz de combustion sont réorientés par un distributeur pour présenter un mouvement giratoire avec une inclinaison supérieure à 70° par rapport à l'axe longitudinal de la turbomachine. Une telle inclinaison est indispensable pour produire l'angle d'attaque nécessaire à la force mécanique d'entraînement en rotation de la roue mobile du premier étage de la section de turbine.The circulation of air through this assembly is generally carried out as follows: the compressed air from the last stage of the compression section has a natural rotational movement with an inclination of the order of 35 ° to 45 ° ° with respect to the longitudinal axis of the turbomachine, tilt which varies according to the speed of the turbomachine (speed of rotation). At its entry into the combustion section, this compressed air is straightened in the longitudinal axis of the turbomachine (that is to say that the inclination of the air with respect to the longitudinal axis of the turbomachine is brought back at 0 °) via an air rectifier. The air in the combustion chamber is then mixed with fuel so as to ensure satisfactory combustion and the gases from this combustion continue a course generally along the longitudinal axis of the turbomachine to reach the turbine section. At the latter, the combustion gases are reoriented by a distributor to present a gyratory movement with an inclination greater than 70 ° with respect to the longitudinal axis of the turbine engine. Such inclination is essential to produce the angle of attack required for the mechanical force driving in rotation of the moving wheel of the first stage of the turbine section.

Une telle distribution angulaire de l'air traversant la turbomachine présente de nombreux inconvénients. En effet, l'air qui sort naturellement du dernier étage de la section de compression avec un angle compris entre 35° et 45° est successivement redressé (angle ramené à 0°) à son entrée dans la section de combustion puis réorienté avec un angle supérieur à 70° à son entrée dans la section de turbine. Ces modifications angulaires successives de la distribution de l'air au travers de la turbomachine nécessitent des efforts aérodynamiques intenses produits par le redresseur de la section de compression et le distributeur de la section de turbine, efforts aérodynamiques qui sont particulièrement préjudiciables pour le rendement global de la turbomachine.Such angular distribution of the air passing through the turbomachine has many disadvantages. Indeed, the air that naturally leaves the last stage of the compression section with an angle between 35 ° and 45 ° is successively rectified (angle reduced to 0 °) at its entry into the combustion section and then reoriented with an angle greater than 70 ° at its entry into the turbine section. These successive angular modifications of the distribution of air through the turbomachine require intense aerodynamic efforts produced by the rectifier of the compression section and the distributor of the turbine section, aerodynamic forces which are particularly detrimental to the overall efficiency of the turbomachine. the turbomachine.

Objet et résumé de l'inventionObject and summary of the invention

La présente invention vise à remédier aux inconvénients précités en proposant une chambre de combustion de turbomachine pouvant être alimentée par un air qui possède un mouvement de rotation par rapport à l'axe longitudinal de la turbomachine.The present invention aims to overcome the aforementioned drawbacks by providing a turbomachine combustion chamber that can be powered by an air that has a rotational movement with respect to the longitudinal axis of the turbomachine.

Ce but est atteint grâce à une chambre de combustion comprenant :

  • une paroi annulaire interne d'axe longitudinal,
  • une paroi annulaire externe centrée sur l'axe longitudinal et entourant la paroi interne de façon à délimiter avec celle-ci un espace annulaire formant un foyer de combustion, et
  • une pluralité de systèmes d'injection de carburant comportant des injecteurs pilote alternant circonférentiellement avec des injecteurs plein gaz,
caractérisée en ce qu'elle comprend en outre au moins une ouverture d'admission d'air débouchant dans le foyer de combustion à l'extrémité amont de celui-ci et selon une direction sensiblement longitudinale ;
en ce que la paroi externe comporte une pluralité de cavités pilote régulièrement réparties autour de l'axe longitudinal, chaque cavité pilote s'étendant longitudinalement entre les deux extrémités longitudinales de la paroi externe et radialement vers l'extérieur de celle-ci, les cavités pilote étant alimentées en air extérieur à la chambre de combustion selon une même direction sensiblement circonférentielle ; et
en ce que chaque injecteur pilote débouche radialement dans une cavité pilote et chaque injecteur plein gaz débouche radialement entre deux cavités pilotes adjacentes.This goal is achieved by a combustion chamber comprising:
  • an inner annular wall of longitudinal axis,
  • an outer annular wall centered on the longitudinal axis and surrounding the inner wall so as to define therewith an annular space forming a combustion chamber, and
  • a plurality of fuel injection systems comprising pilot injectors alternating circumferentially with full-throttle injectors,
characterized in that it further comprises at least one air inlet opening opening into the combustion chamber at the upstream end thereof and in a substantially longitudinal direction;
in that the outer wall comprises a plurality of pilot cavities regularly distributed around the longitudinal axis, each pilot cavity extending longitudinally between the two ends longitudinally of the outer wall and radially outwardly thereof, the pilot cavities being supplied with air outside the combustion chamber in the same substantially circumferential direction; and
in that each pilot injector opens radially into a pilot cavity and each full-gas injector opens radially between two adjacent pilot cavities.

La chambre de combustion selon l'invention peut être alimentée par un air ayant un mouvement de rotation autour de l'axe longitudinal de la turbomachine. L'inclinaison naturelle de l'air en sortie de la section de compression de la turbomachine peut donc être maintenue au travers de la chambre de combustion. Ainsi, l'effort aérodynamique nécessaire à l'entraînement en rotation du premier étage de la section de turbine de la turbomachine est considérablement diminué. Cette forte diminution des efforts aérodynamiques engendre un gain de rendement de la turbomachine. Par ailleurs, le redresseur de la section de compression et le distributeur de la section de turbine peuvent être simplifiés, voire supprimés, ce qui représente un gain de masse et une diminution des coûts de production.The combustion chamber according to the invention can be supplied with air having a rotational movement about the longitudinal axis of the turbomachine. The natural inclination of the air at the outlet of the compression section of the turbomachine can therefore be maintained through the combustion chamber. Thus, the aerodynamic force required for rotating the first stage of the turbine section of the turbomachine is considerably reduced. This sharp decrease in aerodynamic forces generates a gain in efficiency of the turbomachine. In addition, the rectifier of the compression section and the distributor of the turbine section can be simplified or even eliminated, which represents a saving in weight and a reduction in production costs.

En outre, la présence des cavités pilote, qui sont carburées uniquement pour les régimes de ralenti de la turbomachine, permet d'obtenir une stabilisation de la flamme de combustion pour tous les régimes de fonctionnement de la turbomachine.In addition, the presence of pilot cavities, which are carburized only for the idling speeds of the turbomachine, provides a stabilization of the combustion flame for all operating conditions of the turbomachine.

Selon une disposition avantageuse, chaque cavité pilote est fermée à son extrémité amont et ouverte à son extrémité aval.According to an advantageous arrangement, each pilot cavity is closed at its upstream end and open at its downstream end.

Selon une autre disposition avantageuse, chaque cavité pilote est délimitée circonférentiellement par deux cloisons sensiblement radiales, l'une de ces cloisons comportant une pluralité d'orifices d'injection d'air s'ouvrant vers l'extérieur de la chambre de combustion et débouchant dans ladite cavité pilote. De préférence, l'autre cloison de chaque cavité pilote présente, en coupe transversale, une section sensiblement curviligne.According to another advantageous arrangement, each pilot cavity is delimited circumferentially by two substantially radial partitions, one of these partitions comprising a plurality of air injection orifices opening towards the outside of the combustion chamber and opening in said pilot cavity. Preferably, the other partition of each pilot cavity has, in cross section, a substantially curvilinear section.

Selon encore une autre disposition avantageuse, les injecteurs plein gaz sont décalés axialement vers l'aval par rapport aux injecteurs pilote. En effet, la flamme issue des injecteurs pilote a besoin d'un temps de séjour dans le foyer de combustion qui est plus élevé que la flamme issue des injecteurs plein gaz.According to yet another advantageous arrangement, the full-throttle injectors are offset axially downstream relative to the pilot injectors. Indeed, the flame from the pilot injectors needs a residence time in the combustion chamber which is higher than the flame from the injectors full throttle.

La chambre de combustion peut être dépourvue de paroi reliant transversalement les extrémités longitudinales amont des parois interne et externe. L'absence d'une telle paroi (appelée fond de chambre) permet de préserver au maximum la rotation de l'air provenant de la section de compression de la turbomachine.The combustion chamber may be devoid of wall connecting transversely upstream longitudinal ends of the inner and outer walls. The absence of such a wall (called chamber bottom) makes it possible to preserve as much as possible the rotation of the air coming from the compression section of the turbomachine.

Selon encore une autre disposition avantageuse, les systèmes d'injection de carburant sont dépourvus de systèmes d'air associés.According to yet another advantageous arrangement, the fuel injection systems are devoid of associated air systems.

La chambre de combustion peut comprendre en outre un carénage annulaire interne qui est monté sur la paroi interne dans le prolongement amont de celle-ci et un carénage annulaire externe qui est monté sur la paroi externe dans le prolongement amont de celle-ci.The combustion chamber may further comprise an inner annular fairing which is mounted on the inner wall in the upstream extension thereof and an outer annular fairing which is mounted on the outer wall in the upstream extension thereof.

L'invention a également pour objet une turbomachine comprenant une chambre de combustion telle que définie précédemment.The invention also relates to a turbomachine comprising a combustion chamber as defined above.

Brève description des dessinsBrief description of the drawings

D'autres caractéristiques et avantages de la présente invention ressortiront de la description faite ci-dessous, en référence aux dessins annexés qui en illustrent un exemple de réalisation dépourvu de tout caractère limitatif. Sur les figures :

  • la figure 1 est une vue partielle en coupe longitudinale d'une turbomachine aéronautique équipée d'une chambre de combustion selon l'invention ;
  • la figure 2 est une vue en perspective de la chambre de combustion de la figure 1 ;
  • la figure 3 est une vue de face de la chambre de combustion de la figure 2 ;
  • les figures 4 et 5 sont des vues en coupe de la figure 3 respectivement selon IV et V ; et
  • la figure 6 est une vue partielle de face d'une chambre de combustion selon variante de réalisation de l'invention.
Other features and advantages of the present invention will emerge from the description given below, with reference to the accompanying drawings which illustrate an embodiment having no limiting character. In the figures:
  • the figure 1 is a partial view in longitudinal section of an aviation turbine engine equipped with a combustion chamber according to the invention;
  • the figure 2 is a perspective view of the combustion chamber of the figure 1 ;
  • the figure 3 is a front view of the combustion chamber of the figure 2 ;
  • the Figures 4 and 5 are sectional views of the figure 3 respectively according to IV and V; and
  • the figure 6 is a partial front view of a combustion chamber according to an embodiment of the invention.

Description détaillée de modes de réalisationDetailed description of embodiments

La turbomachine partiellement représentée sur la figure 1 possède un axe longitudinal X-X. Selon cet axe, elle comporte notamment : une section annulaire de compression 100, une section annulaire de combustion 200 disposée en sortie de la section de compression 100 selon le sens d'écoulement de l'air traversant la turbomachine, et une section annulaire de turbine 300 disposée en sortie de la section de combustion 200. L'air injecté dans la turbomachine traverse donc successivement la section de compression 100, puis la section de combustion 200 et enfin la section de turbine 300.The turbomachine partially shown on the figure 1 has a longitudinal axis XX. According to this axis, it comprises in particular: an annular compression section 100, an annular combustion section 200 disposed at the outlet of the section of compression 100 according to the direction of flow of the air passing through the turbomachine, and an annular turbine section 300 disposed at the outlet of the combustion section 200. The air injected into the turbomachine therefore passes successively through the compression section 100, then the combustion section 200 and finally the turbine section 300.

La section de compression 100 se présente sous la forme d'une pluralité d'étages de roues mobiles 102 portant chacune des aubes 104 (seul le dernier étage de la section de compression est représenté sur la figure 1). Les aubes 104 de ces étages sont disposées dans un canal annulaire 106 traversé par l'air de la turbomachine et dont la section diminue d'amont en aval. Ainsi, à mesure que l'air injecté dans la turbomachine traverse la section de compression, il est de plus en plus comprimé.The compression section 100 is in the form of a plurality of stages of movable wheels 102 each carrying blades 104 (only the last stage of the compression section is shown in FIG. figure 1 ). The blades 104 of these stages are disposed in an annular channel 106 through which air flows through the turbomachine and whose section decreases from upstream to downstream. Thus, as the air injected into the turbomachine passes through the compression section, it is more and more compressed.

La section de combustion 200 se présente également sous la forme d'un canal annulaire dans lequel l'air comprimé issu de la section de compression 100 est mélangé à du carburant pour y être brûlé. A cet effet, la section de combustion comporte une chambre de combustion 202 à l'intérieur de laquelle est brûlé le mélange air/carburant (cette chambre est détaillée ultérieurement).The combustion section 200 is also in the form of an annular channel in which the compressed air from the compression section 100 is mixed with fuel for burning there. For this purpose, the combustion section comprises a combustion chamber 202 inside which is burned the air / fuel mixture (this chamber is detailed later).

La section de combustion 200 comporte également un carter de turbomachine formé d'une enveloppe annulaire externe 204 centrée sur l'axe longitudinal X-X de la turbomachine et d'une enveloppe annulaire interne 206 qui est fixée de façon coaxiale à l'intérieur de l'enveloppe externe. Un espace annulaire 208 formé entre ces deux enveloppes 204, 206 reçoit de l'air comprimé provenant de la section de compression 100 de la turbomachine.The combustion section 200 also comprises a turbomachine casing formed of an outer annular casing 204 centered on the longitudinal axis XX of the turbomachine and an inner annular casing 206 which is fixed coaxially inside the casing. outer envelope. An annular space 208 formed between these two envelopes 204, 206 receives compressed air from the compression section 100 of the turbomachine.

La section de turbine 300 de la turbomachine est formée par une pluralité d'étages de roues mobiles 302 portant chacune des aubes 304 (seul le premier étage de la section de turbine est représenté sur la figure 1). Les aubes 304 de ces étages sont disposées dans un canal annulaire 306 traversé par les gaz issus de la section de combustion 200.The turbomachine section 300 of the turbomachine is formed by a plurality of stages of movable wheels 302 each carrying blades 304 (only the first stage of the turbine section is shown in FIG. figure 1 ). The blades 304 of these stages are arranged in an annular channel 306 traversed by the gases coming from the combustion section 200.

En entrée du premier étage 302 de la section de turbine 300, les gaz issus de la section de combustion doivent présenter une inclinaison par rapport à l'axe longitudinal X-X de la turbomachine qui soit suffisante pour entraîner en rotation les différents étages de la section de turbine.At the inlet of the first stage 302 of the turbine section 300, the gases coming from the combustion section must have an inclination relative to the longitudinal axis XX of the turbomachine which is sufficient to rotate the different stages of the turbine section. turbine.

A cet effet, un distributeur 308 est monté directement en aval de la chambre de combustion 202 et en amont du premier étage 302 de la section de turbine 300. Ce distributeur 308 se compose d'une pluralité d'aubes radiales fixes 310 dont l'inclinaison par rapport à l'axe longitudinal X-X de la turbomachine permet de donner aux gaz issus de la section de combustion 200 l'inclinaison nécessaire à l'entraînement en rotation des différents étages de la section de turbine.For this purpose, a distributor 308 is mounted directly downstream of the combustion chamber 202 and upstream of the first stage 302 of the turbine section 300. This distributor 308 consists of a plurality of fixed radial vanes 310 of which inclination with respect to the longitudinal axis XX of the turbomachine makes it possible to give the gases coming from the combustion section 200 the inclination necessary for driving in rotation the different stages of the turbine section.

Dans les turbomachines classiques, la distribution de l'air traversant successivement la section de compression 100, la section de combustion 200 et la section de turbine 300 s'opère de la façon suivante. L'air comprimé issu du dernier étage 102 de la section de compression 100 possède naturellement un mouvement giratoire avec une inclinaison de l'ordre de 35° à 45° par rapport à l'axe longitudinal X-X de la turbomachine. Par l'intermédiaire du redresseur d'air 210 de la section de combustion 200, cet angle d'inclinaison est ramené à 0°. Enfin, au niveau de l'entrée de la section de turbine 300, les gaz issus de la combustion sont réorientés par les aubes fixes 310 du distributeur 308 de cette dernière pour leur donner un mouvement giratoire avec une inclinaison par rapport à l'axe longitudinal X-X qui est supérieure à 70°.In conventional turbomachines, the distribution of the air successively passing through the compression section 100, the combustion section 200 and the turbine section 300 takes place as follows. The compressed air from the last stage 102 of the compression section 100 naturally has a gyratory movement with an inclination of the order of 35 ° to 45 ° relative to the longitudinal axis X-X of the turbomachine. Through the air rectifier 210 of the combustion section 200, this inclination angle is reduced to 0 °. Finally, at the inlet of the turbine section 300, the gases resulting from the combustion are redirected by the blades 310 of the distributor 308 of the latter to give them a gyratory movement with an inclination with respect to the longitudinal axis XX which is greater than 70 °.

Selon l'invention, il est prévu une nouvelle architecture de la chambre de combustion 202 qui peut être alimentée par un air possédant un mouvement de rotation autour de l'axe longitudinal X-X de la turbomachine. Grâce à une telle architecture, il est possible de conserver l'inclinaison naturelle de l'air comprimé issu du dernier étage de la section de compression sans avoir à le redresser dans l'axe longitudinal X-X. De même, il n'est plus nécessaire que les aubes fixes 310 du distributeur 308 de la section de turbine 300 présentent une inclinaison aussi importante pour produire l'angle d'attaque nécessaire à la force mécanique d'entraînement en rotation de la roue mobile 302 du premier étage de la section de turbine.According to the invention, there is provided a new architecture of the combustion chamber 202 which can be powered by an air having a rotational movement about the longitudinal axis X-X of the turbomachine. With such an architecture, it is possible to maintain the natural inclination of the compressed air from the last stage of the compression section without having to straighten it in the longitudinal axis X-X. Likewise, it is no longer necessary for the stationary blades 310 of the distributor 308 of the turbine section 300 to have such a large inclination to produce the angle of attack required for the mechanical driving force in rotation of the moving wheel. 302 of the first stage of the turbine section.

A cet effet, la chambre de combustion 202 selon l'invention comprend une paroi annulaire interne 212 centrée sur l'axe longitudinal X-X de la turbomachine, et une paroi annulaire externe 214 également centrée sur l'axe longitudinal X-X et entourant la paroi interne de façon à délimiter avec celle-ci un espace annulaire 216 formant un foyer de combustion.For this purpose, the combustion chamber 202 according to the invention comprises an inner annular wall 212 centered on the longitudinal axis XX of the turbomachine, and an outer annular wall 214 also centered on the longitudinal axis XX and surrounding the inner wall of the engine. to define with it an annular space 216 forming a combustion chamber.

La chambre de combustion 202 selon l'invention comprend en outre au moins une ouverture d'admission d'air 218 qui débouche dans le foyer de combustion 216 à l'extrémité amont de celui-ci et selon une direction sensiblement longitudinale. La section de cette ouverture d'admission d'air est adaptée pour assurer le fonctionnement du foyer de combustion.The combustion chamber 202 according to the invention further comprises at least one air inlet opening 218 which opens into the combustion chamber 216 at the upstream end thereof and in a substantially longitudinal direction. The section of this air intake opening is adapted to ensure the operation of the combustion chamber.

Plus précisément, comme représenté sur la figure 1, la chambre de combustion étant dépourvue de paroi (appelée fond de chambre) reliant transversalement les extrémités longitudinales amont des parois interne et externe, cette ouverture d'admission d'air 218 est formée entre les extrémités amont des parois interne 212 et externe 214 de la chambre de combustion.More precisely, as represented on the figure 1 the combustion chamber having no wall (called chamber bottom) connecting transversely the longitudinal ends upstream of the inner and outer walls, this air inlet opening 218 is formed between the upstream ends of the inner and outer walls 212 and 214 of the combustion chamber.

La chambre de combustion 202 selon l'invention comprend encore une pluralité de systèmes d'injection de carburant 220 répartis sur la paroi externe 214 autour de l'axe longitudinal X-X de la turbomachine et débouchant dans le foyer de combustion 216 selon une direction sensiblement radiale.The combustion chamber 202 according to the invention also comprises a plurality of fuel injection systems 220 distributed on the outer wall 214 around the longitudinal axis XX of the turbomachine and opening into the combustion chamber 216 in a substantially radial direction .

Comme représenté sur les figures 2 et 3, les systèmes d'injection de carburant 220 comportent des injecteurs pilote 220a alternant circonférentiellement avec des injecteurs plein gaz 220b, les injecteurs plein gaz étant de préférence décalés axialement vers l'aval par rapport aux injecteurs pilote.As shown on figures 2 and 3 , the fuel injection systems 220 comprise pilot injectors 220a circumferentially alternating with full-throttle injectors 220b, the full-throttle injectors preferably being offset axially downstream relative to the pilot injectors.

Classiquement, les injecteurs pilote 220a assurent l'allumage et les phases de ralenti de la turbomachine et les injecteurs plein gaz 220b interviennent dans les phases de décollage, de montée et de croisière. En général, les injecteurs pilote sont alimentés en carburant en permanence tandis que les injecteurs plein gaz ne sont alimentés qu'au-delà d'un certain régime déterminé.Conventionally, the pilot injectors 220a provide ignition and idle phases of the turbomachine and the 220b full-throttle injectors are involved in the take-off, climb and cruise phases. In general, the pilot injectors are fueled continuously while the full-throttle injectors are only fed beyond a certain determined speed.

Selon une caractéristique particulière avantageuse de l'invention, les systèmes d'injection de carburant 220 sont dépourvus de systèmes d'air associés tels que des vrilles d'air qui permettent, de façon connue en soi, de générer un écoulement d'air rotatif à l'intérieur du foyer de combustion dans le but de stabiliser la flamme de combustion.According to one particular advantageous characteristic of the invention, the fuel injection systems 220 are devoid of associated air systems such as air swirlers which make it possible, in a manner known per se, to generate a rotary air flow. inside the combustion chamber in order to stabilize the combustion flame.

Ainsi, les injecteurs pilote et plein gaz de la chambre de combustion sont de conception très simple et de fonctionnement très fiable puisqu'ils sont réduits à leur plus simple fonction, à savoir injecter du carburant. De plus, les injecteurs pilote 220a sont du même type que les injecteurs plein gaz 220b.Thus, the pilot and full throttle injectors of the combustion chamber are of very simple design and very reliable operation since they are reduced to their simplest function, namely to inject fuel. In addition, the pilot injectors 220a are of the same type as the full throttle injectors 220b.

Toujours selon l'invention, la paroi externe 214 de la chambre de combustion comporte une pluralité de cavités pilote 222 qui sont régulièrement réparties autour de l'axe longitudinal X-X.Still according to the invention, the outer wall 214 of the combustion chamber comprises a plurality of pilot cavities 222 which are regularly distributed around the longitudinal axis X-X.

Comme représenté sur la figure 2, chaque cavité pilote 222 s'étend, d'une part longitudinalement entre les deux extrémités longitudinales (amont et aval) de la paroi externe 214, et d'autre part radialement vers l'extérieur de celle-ci. En d'autres termes, la paroi externe 214 est profilée avec une pluralité de cavités 222 faisant saillie vers l'extérieur de la paroi.As shown on the figure 2 each pilot cavity 222 extends, firstly longitudinally between the two longitudinal ends (upstream and downstream) of the outer wall 214, and secondly radially outwardly thereof. In other words, the outer wall 214 is profiled with a plurality of cavities 222 protruding outwardly from the wall.

De façon plus précise, les cavités pilote 222 sont chacune délimitées circonférentiellement par deux cloisons 224 qui font chacune saillies radialement vers l'extérieur par rapport à la paroi externe 214. Comme représenté sur les figures 2 et 5, l'une de ces cloisons présente une pluralité d'orifices d'injection d'air 226 qui permettent d'injecter de l'air extérieur à la chambre de combustion dans la cavité pilote selon une direction circonférentielle.More specifically, the pilot cavities 222 are each delimited circumferentially by two partitions 224 which each project radially outwardly with respect to the outer wall 214. As shown in FIGS. figures 2 and 5 , one of these partitions has a plurality of air injection orifices 226 which make it possible to inject air outside the combustion chamber into the pilot cavity in a circumferential direction.

Il est à noter que l'injection circonférentielle d'air est réalisée selon un même sens de rotation (celui des aiguilles d'une montre pour l'exemple de réalisation des figures 2 et 3) pour l'ensemble des cavités pilote 222 de la chambre de combustion. Par ailleurs, le sens de rotation pour l'injection circonférentielle d'air dans ces cavités pilote est celui de l'air comprimé provenant de la section de compression de la turbomachine.It should be noted that the circumferential injection of air is performed in the same direction of rotation (that of the needles of a watch for the exemplary embodiment of figures 2 and 3 ) for all the pilot cavities 222 of the combustion chamber. Moreover, the direction of rotation for the circumferential injection of air into these pilot cavities is that of the compressed air coming from the compression section of the turbomachine.

Les cavités pilote 222 sont alimentées en carburant par l'intermédiaire des injecteurs pilotes 220a qui débouchent chacun radialement dans l'une de ces cavités. Quant aux injecteurs plein gaz 220b, ils débouchent chacun radialement dans le foyer de combustion entre deux cavités pilote adjacentes.The pilot cavities 222 are supplied with fuel via pilot injectors 220a, each of which opens radially into one of these cavities. As for the full-throttle injectors 220b, they each open radially into the combustion chamber between two adjacent pilot cavities.

Chaque cavité pilote 222 est de préférence fermée à son extrémité amont par une cloison radiale 228 et ouverte à son extrémité aval (voir notamment les figures 2 et 5). Ainsi, l'air qui pénètre dans le foyer de combustion 216 par son ouverture d'admission d'air 218 ne vient pas perturber l'écoulement d'air introduit dans les cavités pilote 222 par les orifices d'injection d'air 226.Each pilot cavity 222 is preferably closed at its upstream end by a radial partition 228 and open at its downstream end (see especially the figures 2 and 5 ). Thus, the air entering the combustion chamber 216 through its air inlet opening 218 does not disturb the flow of air introduced into the pilot cavities 222 by the air injection orifices 226.

Le fonctionnement de la chambre de combustion est le suivant : l'air comprimé provenant de la section de compression 100 et qui est en rotation autour de l'axe longitudinal X-X pénètre dans la section de combustion 200. Cet air se répartit en deux écoulements : un écoulement « interne » et un écoulement « externe ». L'écoulement externe contourne la chambre de combustion 202 et alimente les cavités pilote 222 après avoir refroidi la paroi externe 214 de la chambre de combustion et le carter externe 204 de la section de combustion. Cet air externe est injecté dans ces cavités pilote par l'intermédiaire des orifices d'injection d'air 226 selon le sens de rotation de l'air à son entrée dans la section de combustion. Dans ces cavités pilote, l'air est mélangé et brûlé au carburant injecté par les injecteurs pilote 220a. Quant à l'écoulement interne qui représente l'écoulement principal, il pénètre dans le foyer de combustion 216 par l'ouverture d'admission d'air 218 pour être mélangé et brûlé au carburant injecté par les injecteurs plein gaz 220b. La stabilisation de la flamme de combustion est obtenue grâce à la « carburation » des cavités pilote.The operation of the combustion chamber is as follows: the compressed air coming from the compression section 100 and which is rotated about the longitudinal axis XX enters the combustion section 200. This air is divided into two flows: "internal" flow and "external" flow. The external flow bypasses the combustion chamber 202 and feeds the pilot cavities 222 after cooling the outer wall 214 of the combustion chamber and the outer casing 204 of the combustion section. This external air is injected into these pilot cavities via the air injection orifices 226 in the direction of rotation of the air at its entry into the combustion section. In these pilot cavities, the air is mixed and burnt with the fuel injected by the pilot injectors 220a. As for the internal flow that represents the main flow, it enters the combustion chamber 216 through the air inlet opening 218 to be mixed and burned fuel injected by the full-throttle injectors 220b. Stabilization of the combustion flame is obtained thanks to the "carburation" of the pilot cavities.

On décrira maintenant des variantes de réalisation de la chambre de combustion selon l'invention.We will now describe embodiments of the combustion chamber according to the invention.

Dans l'exemple de réalisation des figures 2 et 3, la cloison longitudinale 224 de chaque cavité pilote 222 qui n'est pas munie d'orifices d'injection d'air présente, en coupe transversale, une section sensiblement curviligne (contrairement à l'autre paroi qui est sensiblement plane). La courbure de ces parois permet d'accompagner le mouvement de rotation de l'air injecté dans les cavités pilote par les orifices d'injection d'air 226.In the exemplary embodiment of figures 2 and 3 , the longitudinal partition 224 of each pilot cavity 222 which is not provided with air injection orifices has, in cross-section, a substantially curvilinear section (unlike the other wall which is substantially flat). The curvature of these walls makes it possible to accompany the rotational movement of the air injected into the pilot cavities by the air injection orifices 226.

Au contraire, dans la variante de réalisation de la figure 6, les deux cloisons longitudinales 224 délimitant circonférentiellement chaque cavité pilote 222 sont sensiblement planes et s'étendent chacune selon une direction radiale.On the contrary, in the variant embodiment of the figure 6 , the two longitudinal partitions 224 circumferentially delimiting each pilot cavity 222 are substantially flat and each extend in a radial direction.

De manière générale, le nombre et les dimensions géométriques des cavités pilote 222 de la chambre de combustion peuvent varier en fonction des besoins. Il en est de même du nombre, des dimensions et du positionnement des orifices d'injection d'air 226 dans ces cavités.In general, the number and the geometric dimensions of the pilot cavities 222 of the combustion chamber may vary according to the needs. The same is true of the number, the dimensions and the positioning of the air injection orifices 226 in these cavities.

Comme représenté à la figure 1, la chambre de combustion 202 peut également comporter un carénage annulaire interne 230 qui est monté sur la paroi interne 212 dans le prolongement amont de celle-ci et un carénage annulaire externe 232 qui est monté sur la paroi externe 214 dans le prolongement amont de celle-ci. La présence de ces carénages 230, 232 permet de régler le débit d'air entrant dans la chambre de combustion 202 et celui la contournant.As represented in figure 1 , the combustion chamber 202 may also comprise an internal annular fairing 230 which is mounted on the inner wall 212 in the upstream extension thereof and an outer annular fairing 232 which is mounted on the outer wall 214 in the upstream extension thereof. The presence of these shrouds 230, 232 makes it possible to regulate the flow rate of air entering the combustion chamber 202 and that bypassing it.

Enfin, la paroi externe 214 de la chambre de combustion peut comporter à son extrémité aval une bride annulaire 234 s'étendant radialement vers l'extérieur de la paroi, cette bride étant munie d'une pluralité de trous 236 régulièrement répartis autour de l'axe longitudinal X-X et destinés à alimenter en air de refroidissement la section de turbine 300.Finally, the outer wall 214 of the combustion chamber may comprise at its downstream end an annular flange 234 extending radially outwardly of the wall, this flange being provided with a plurality of holes 236 regularly distributed around the longitudinal axis XX and intended to supply cooling air to the turbine section 300.

Claims (9)

Chambre de combustion (202) de turbomachine comprenant : une paroi annulaire interne (212) d'axe longitudinal (X-X), une paroi annulaire externe (214) centrée sur l'axe longitudinal et entourant la paroi interne de façon à délimiter avec celle-ci un espace annulaire (216) formant un foyer de combustion, et une pluralité de systèmes d'injection de carburant (220) comportant des injecteurs pilote (220a) alternant circonférentiellement avec des injecteurs plein gaz (220b), caractérisée en ce qu'elle comprend en outre au moins une ouverture d'admission d'air (218) débouchant dans le foyer de combustion à l'extrémité amont de celui-ci et selon une direction sensiblement longitudinale ;
en ce que la paroi externe (214) comporte une pluralité de cavités pilote (222) régulièrement réparties autour de l'axe longitudinal, chaque cavité pilote s'étendant longitudinalement entre les deux extrémités longitudinales de la paroi externe et radialement vers l'extérieur de celle-ci, les cavités pilote étant alimentées en air extérieur à la chambre de combustion selon une même direction sensiblement circonférentielle ; et
en ce que chaque injecteur pilote (220a) débouche radialement dans une cavité pilote (222) et chaque injecteur plein gaz (220b) débouche radialement entre deux cavités pilotes adjacentes.A turbomachine combustion chamber (202) comprising: an inner annular wall (212) of longitudinal axis (XX), an outer annular wall (214) centered on the longitudinal axis and surrounding the inner wall so as to define therewith an annular space (216) forming a combustion chamber, and a plurality of fuel injection systems (220) having pilot injectors (220a) alternating circumferentially with full-throttle injectors (220b), characterized in that it further comprises at least one air inlet opening (218) opening into the combustion chamber at the upstream end thereof and in a substantially longitudinal direction;
in that the outer wall (214) comprises a plurality of pilot cavities (222) regularly distributed around the longitudinal axis, each pilot cavity extending longitudinally between the two longitudinal ends of the outer wall and radially outwardly of the latter, the pilot cavities being supplied with air outside the combustion chamber in the same substantially circumferential direction; and
in that each pilot injector (220a) opens radially into a pilot cavity (222) and each full-gas injector (220b) opens radially between two adjacent pilot cavities. Chambre de combustion selon la revendication 1, dans laquelle chaque cavité pilote (222) est fermée à son extrémité amont et ouverte à son extrémité aval.Combustion chamber according to claim 1, wherein each pilot cavity (222) is closed at its upstream end and open at its downstream end. Chambre de combustion selon l'une des revendications 1 et 2, dans laquelle chaque cavité pilote (222) est délimitée circonférentiellement par deux cloisons (224) sensiblement radiales, l'une de ces cloisons comportant une pluralité d'orifices d'injection d'air (226) s'ouvrant vers l'extérieur de la chambre de combustion et débouchant dans ladite cavité pilote.Combustion chamber according to one of claims 1 and 2, wherein each pilot cavity (222) is delimited circumferentially by two substantially radial partitions (224), one of these partitions comprising a plurality of injection orifices of air (226) opening towards the outside of the combustion chamber and opening into said pilot cavity. Chambre de combustion selon la revendication 3, dans laquelle l'autre cloison de chaque cavité pilote (222) présente, en coupe transversale, une section sensiblement curviligne.Combustion chamber according to claim 3, wherein the other partition of each pilot cavity (222) has, in cross-section, a substantially curvilinear section. Chambre de combustion selon l'une quelconque des revendications 1 à 4, dans laquelle les injecteurs plein gaz (220b) sont décalés axialement vers l'aval par rapport aux injecteurs pilote (220a).Combustion chamber according to any one of claims 1 to 4, wherein the full-throttle injectors (220b) are offset axially downstream relative to the pilot injectors (220a). Chambre de combustion selon l'une quelconque des revendications 1 à 5, dans laquelle elle est dépourvue de paroi reliant transversalement les extrémités longitudinales amont des parois interne (212) et externe (214).Combustion chamber according to any one of claims 1 to 5, wherein it has no wall connecting transversely upstream longitudinal ends of the inner walls (212) and outer (214). Chambre de combustion selon l'une quelconque des revendications 1 à 6, dans laquelle les systèmes d'injection de carburant (220) sont dépourvus de systèmes d'air associés.Combustion chamber according to any one of claims 1 to 6, wherein the fuel injection systems (220) are devoid of associated air systems. Chambre de combustion selon l'une quelconque des revendications 1 à 7, comportant en outre un carénage annulaire interne (230) qui est monté sur la paroi interne (212) dans le prolongement amont de celle-ci et un carénage annulaire externe (232) qui est monté sur la paroi externe (214) dans le prolongement amont de celle-ci.Combustion chamber according to any one of claims 1 to 7, further comprising an inner annular fairing (230) which is mounted on the inner wall (212) in the upstream extension thereof and an outer annular fairing (232) which is mounted on the outer wall (214) in the upstream extension thereof. Turbomachine caractérisée en ce qu'elle comporte une chambre de combustion (202) selon l'une quelconque des revendications 1 à 8.Turbomachine characterized in that it comprises a combustion chamber (202) according to any one of claims 1 to 8.
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FR2917487B1 (en) 2009-10-02
US7673456B2 (en) 2010-03-09
RU2008124152A (en) 2009-12-20
JP5084626B2 (en) 2012-11-28
CN101324344A (en) 2008-12-17
IL192052A (en) 2011-07-31
CN101324344B (en) 2011-08-17
RU2478880C2 (en) 2013-04-10
EP2003399A3 (en) 2013-07-31
CA2634615A1 (en) 2008-12-14
EP2003399B1 (en) 2014-04-30
US20080307792A1 (en) 2008-12-18
CA2634615C (en) 2014-08-05
FR2917487A1 (en) 2008-12-19
JP2008309466A (en) 2008-12-25
IL192052A0 (en) 2009-02-11

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