EP2034245B1 - Gas turbine combustion chamber with helicoidal air circulation - Google Patents

Gas turbine combustion chamber with helicoidal air circulation Download PDF

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Publication number
EP2034245B1
EP2034245B1 EP08163522A EP08163522A EP2034245B1 EP 2034245 B1 EP2034245 B1 EP 2034245B1 EP 08163522 A EP08163522 A EP 08163522A EP 08163522 A EP08163522 A EP 08163522A EP 2034245 B1 EP2034245 B1 EP 2034245B1
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EP
European Patent Office
Prior art keywords
combustion chamber
wall
air
turbomachine
longitudinal axis
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EP08163522A
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German (de)
French (fr)
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EP2034245A1 (en
Inventor
Michel Pierre Cazalens
Romain Nicolas Lunel
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Safran Aircraft Engines SAS
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SNECMA SAS
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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
    • F23R3/58—Cyclone or vortex type combustion chambers
    • 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 a satisfactory combustion and the gases resulting 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 ° relative to the longitudinal axis of the turbomachine.
  • 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 forces 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.
  • a turbomachine combustion chamber according to the preamble of claim 1 is known from the document U.S. 5,025,622 A .
  • 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 is supplied with air via the internal and external cavities in a substantially circumferential direction.
  • the combustion chamber according to the invention can thus 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 design of the high-pressure turbine distributor can be simplified and the aerodynamic force required to bring the flow in the axis of the turbomachine substantially decreased. 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 being simplified, this can lead to a saving in weight and a reduction in production costs.
  • certain internal and external steps comprise a substantially radial wall provided with a plurality of air injection orifices opening towards the outside of the combustion chamber and opening into the adjacent internal or external cavity.
  • the internal and external steps comprise another wall which has, in cross section, a substantially curvilinear section.
  • the fuel injection systems comprise pilot injectors alternating circumferentially with full-throttle injectors.
  • the full-throttle injectors are preferably axially offset downstream relative to the pilot injectors.
  • 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 fuel injection systems are devoid of associated air systems (which generally allow the air to be rotated so as to create a recirculation in order to stabilize the combustion flame ).
  • 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. Along this axis, it comprises in particular an annular compression section 100, an annular combustion section 200 disposed at the outlet of the compression section 100 in the direction of flow of the air passing through the turbomachine, and a section annular turbine 300 disposed at the output of the combustion section 200.
  • the air injected into the turbomachine thus successively passes 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 whose inclination relative 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 rotating 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, an outer annular wall 214 also centered on the longitudinal axis XX and surrounding the inner wall of to define therewith an annular space 216 forming a combustion focus, and a transverse annular wall 218 (called chamber bottom) transversely connecting the longitudinal ends upstream of the inner and outer walls.
  • the internal wall 212 of the combustion chamber comprises a plurality of internal steps (or steps) 220 which are regularly distributed around the longitudinal axis X-X. Each of these internal steps 220 extends, firstly longitudinally between the two longitudinal ends (upstream and downstream) of the inner wall, and secondly radially outwardly thereof.
  • the inner surface of the inner wall 212 is profiled with a plurality of steps 220 protruding outwardly from the wall. Furthermore, internal cavity 222 designates the circumferential spacing that is defined between two adjacent internal steps 220.
  • the outer wall 214 of the combustion chamber comprises a plurality of steps (or steps) external 224 evenly distributed around the longitudinal axis X-X.
  • Each external step 224 extends, firstly longitudinally between the two longitudinal ends of the outer wall, and secondly radially inwardly thereof.
  • the outer surface of the outer wall 214 is profiled with a plurality of steps 224 projecting inwardly from the wall.
  • External cavity 226 denotes the circumferential spacing that is defined between two adjacent external steps 224.
  • some of the internal cavities 222 and some of the external cavities 226 are supplied with fuel in a substantially radial direction.
  • the combustion chamber 202 also comprises a plurality of fuel injection systems 228 distributed on the inner walls 212 and outer 214 around the longitudinal axis XX of the turbomachine and opening into the combustion chamber. combustion 216 in a substantially radial direction.
  • the fuel injection systems 228 radially open into some of the internal cavities 222 and some of the outer cavities 226.
  • the fuel injection systems 228 open into all the external cavities 226 and into only one internal cavity 222 out of two.
  • all the internal cavities and all the external cavities can be supplied with fuel; only one external cavity out of two and all internal cavities are fueled; etc.
  • the principle governing the choice of the supply configuration of these cavities is to optimize the performance of the combustion chamber for each point of the flight envelope.
  • fuel injection systems 228 comprise pilot injectors 228a alternating circumferentially with full-throttle injectors 228b.
  • the fuel injection systems 228 supplying the external cavities 226 do indeed comprise an alternation of pilot injectors 228a with full-throttle injectors, and the fuel injection systems 228 supplying the internal cavities 222 comprise full-throttle injectors and injectors. pilot injectors.
  • the pilot injectors 228a provide ignition and idle phases of the turbomachine and the full-throttle injectors 228b are involved in the take-off, climb and cruise phases.
  • the pilot injectors are fueled continuously while the takeoff injectors are only fed beyond a certain determined regime.
  • the fuel injection systems 228 do not have 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.
  • pilot injectors 228a are of the same type as the full-throttle injectors 228b.
  • the full-throttle injectors 228b can be axially offset downstream relative to the pilot injectors 228a.
  • At least some of the internal cavities 222 and some of the external cavities 226 are supplied with air outside the combustion chamber 202 in a same substantially circumferential direction.
  • the internal cavities 222 and external 226 which are supplied with air by means of a plurality of air injection orifices 230 formed in a substantially radial wall 232 of the internal 220 and external steps 224 corresponding.
  • These air injection orifices 230 open towards the outside of the combustion chamber 202 and open into the corresponding internal or external cavity in a substantially circumferential direction.
  • the circumferential injection of air into the combustion chamber 216 is made in the same direction of rotation (that of the needles of a watch for the embodiment of figures 2 and 3 ) for all internal cavities 222 and external 226 of the combustion chamber. Moreover, the direction of rotation for the circumferential injection of air into these cavities is that of the compressed air coming from the compression section of the turbomachine.
  • the air supply of the combustion chamber 206 is only achieved by means of the air injection orifices 230 opening into some of the internal and external cavities in a circumferential direction (a very small proportion of air also enters the combustion chamber through multiperforation holes made in the walls 212, 214 and 218 of the combustion chamber for cooling these walls, these holes are not shown in the figures).
  • the internal and external cavities that are supplied with fuel are not necessarily homogeneous with respect to their radial dimension (that is to say the height of the corresponding step) and circumferential so as to be able to vary the residence time following the cavity considered.
  • the height of the steps is not necessarily constant over the entire length of the wall (that is to say between its upstream and downstream ends).
  • the air flow supplying these cavities may vary depending on the cavity considered.
  • the operation of the combustion chamber is as follows: the compressed air coming from the compression section 100 and rotating about the longitudinal axis XX enters the combustion section 200. This air bypasses the combustion chamber 202 and supplies at least some of the internal and external cavities 222 and 226 after cooling the walls and casings of the combustion chamber. This air is injected into these cavities via the air injection orifices 230 in the direction of rotation of the air at its entry into the combustion section. In some of these air-fed cavities, the air is mixed and burned with the fuel injected by the fuel injection systems 228.
  • the internal and external steps 224 224 of the combustion chamber comprise another wall 232 '(opposite to that 232 provided with air injection orifices) which extends in a substantially circumferential direction and which has, in section transverse, a substantially curvilinear section (unlike the wall 232 which is substantially flat and radial).
  • the curvature of this wall makes it possible to form a ramp to accompany the rotational movement of the air injected into the cavities via the air injection orifices 230.
  • any other wall form (planar or curvilinear) can be envisaged. .
  • the number and the geometrical dimensions of the internal and external cavities 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 in these cavities, as well as the relative circumferential position of the fuel injection systems with respect to the internal and external steps.
  • the inner wall 212 and the outer wall 214 of the combustion chamber may each comprise at their downstream end an annular flange, respectively 234 and 236, which is provided with a plurality of holes 238 regularly distributed about the longitudinal axis XX and for supplying 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)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

Arrière-plan de l'inventionBackground of the invention

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 a satisfactory combustion and the gases resulting 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 ° relative to the longitudinal axis of the turbomachine. 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 forces 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.

Une chambre de combustion de turbomachine selon le préambule de la revendicatiion 1 est connue du document US-5 025 622 A .A turbomachine combustion chamber according to the preamble of claim 1 is known from the document U.S. 5,025,622 A .

Objet et résumé de I'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,
  • une paroi annulaire transversale reliant transversalement les extrémités longitudinales amont des parois interne et externe, et
  • une pluralité de systèmes d'injection de carburant,
  • la paroi interne comportant une pluralité de marches internes régulièrement réparties autour de l'axe longitudinal, chaque marche interne s'étendant longitudinalement entre les deux extrémités longitudinales de la paroi interne et radialement vers l'extérieur de celle-ci, l'espacement circonférentiel entre deux marches internes adjacentes définissant une cavité interne ;
  • la paroi externe comportant une pluralité de marches externes régulièrement réparties autour de l'axe longitudinal, chaque marche externe s'étendant longitudinalement entre les deux extrémités longitudinales de la paroi externe et radialement vers l'intérieur de celle-ci, l'espacement circonférentiel entre deux marches internes adjacentes définissant une cavité externe ; et
  • certaines cavités internes et externes sont alimentées en air extérieur à la chambre de combustion selon une même direction sensiblement circonférentielle caractérisée en ce que ces certaines cavités internes et externes sont également alimentées en carburant selon une direction sensiblement radiale.
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,
  • a transverse annular wall connecting transversely the longitudinal ends upstream of the inner and outer walls, and
  • a plurality of fuel injection systems,
  • the inner wall having a plurality of internal steps evenly distributed about the longitudinal axis, each inner step extending longitudinally between the two ends longitudinally of the inner wall and radially outwardly thereof, the circumferential spacing between two adjacent internal steps defining an internal cavity;
  • the outer wall having a plurality of external steps evenly distributed about the longitudinal axis, each outer step extending longitudinally between the two longitudinal ends of the outer wall and radially inwardly thereof, the circumferential spacing between two adjacent internal steps defining an external cavity; and
  • certain internal and external cavities are supplied with air outside the combustion chamber in the same substantially circumferential direction, characterized in that these internal and external cavities are also supplied with fuel in a substantially radial direction.

L'alimentation en air du foyer de combustion s'effectue par l'intermédiaire des cavités internes et externes selon une direction sensiblement circonférentielle. La chambre de combustion selon l'invention peut ainsi 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. De la sorte, la conception aérodynamique du distributeur de turbine haute-pression peut être simplifiée et l'effort aérodynamique nécessaire pour remettre l'écoulement dans l'axe de la turbomachine sensiblement 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 étant simplifiés, cela peut engendrer un gain de masse et une diminution des coûts de production.The combustion chamber is supplied with air via the internal and external cavities in a substantially circumferential direction. The combustion chamber according to the invention can thus 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. In this way, the aerodynamic design of the high-pressure turbine distributor can be simplified and the aerodynamic force required to bring the flow in the axis of the turbomachine substantially decreased. This sharp decrease in aerodynamic forces generates a gain in efficiency of the turbomachine. Furthermore, the rectifier of the compression section and the distributor of the turbine section being simplified, this can lead to a saving in weight and a reduction in production costs.

En outre, la présence de cavités externes et internes, qui peuvent être 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 external and internal cavities, which can be carburized only for the idling speeds of the turbomachine, provides a stabilization of the combustion flame for all operating modes of the turbomachine.

Selon une disposition avantageuse, certaines marches internes et externes comportent une paroi sensiblement radiale munie d'une pluralité d'orifices d'injection d'air s'ouvrant vers l'extérieur de la chambre de combustion et débouchant dans la cavité interne ou externe adjacente.According to an advantageous arrangement, certain internal and external steps comprise a substantially radial wall provided with a plurality of air injection orifices opening towards the outside of the combustion chamber and opening into the adjacent internal or external cavity. .

Selon une autre disposition avantageuse, les marches internes et externes comportent une autre paroi qui présente, en coupe transversale, une section sensiblement curviligne.According to another advantageous arrangement, the internal and external steps comprise another wall which has, in cross section, a substantially curvilinear section.

Selon encore une autre disposition avantageuse, les systèmes d'injection de carburant comportent des injecteurs pilote alternant circonférentiellement avec des injecteurs plein gaz. Dans ce cas, les injecteurs plein gaz sont de préférence 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 fuel injection systems comprise pilot injectors alternating circumferentially with full-throttle injectors. In this case, the full-throttle injectors are preferably axially offset 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.

Selon encore une autre disposition avantageuse, les systèmes d'injection de carburant sont dépourvus de systèmes d'air associés (qui permettent généralement de mettre l'air en rotation de manière à créer une re-circulation dans le but de stabiliser la flamme de combustion).According to yet another advantageous arrangement, the fuel injection systems are devoid of associated air systems (which generally allow the air to be rotated so as to create a recirculation in order to stabilize the combustion flame ).

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 un mode de réalisation de 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 figure 2 ; et
  • la figure 4 est une vue en coupe selon IV-IV de la figure 3.
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 one embodiment of 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 figure 2 ; and
  • the figure 4 is a sectional view along IV-IV of the figure 3 .

Description détaillée d'un mode de réalisationDetailed description of an embodiment

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. Along this axis, it comprises in particular an annular compression section 100, an annular combustion section 200 disposed at the outlet of the compression section 100 in the direction of flow of the air passing through the turbomachine, and a section annular turbine 300 disposed at the output of the combustion section 200. The air injected into the turbomachine thus successively passes 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 whose inclination relative 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 rotating 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, 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, et une paroi annulaire transversale 218 (appelée fond de chambre) reliant transversalement les extrémités longitudinales amont des parois interne et externe.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, an outer annular wall 214 also centered on the longitudinal axis XX and surrounding the inner wall of to define therewith an annular space 216 forming a combustion focus, and a transverse annular wall 218 (called chamber bottom) transversely connecting the longitudinal ends upstream of the inner and outer walls.

La paroi interne 212 de la chambre de combustion comporte une pluralité de marches (ou marches) internes 220 qui sont régulièrement réparties autour de l'axe longitudinal X-X. Chacune de ces marches internes 220 s'étend, d'une part longitudinalement entre les deux extrémités longitudinales (amont et aval) de la paroi interne, et d'autre part radialement vers l'extérieur de celle-ci.The internal wall 212 of the combustion chamber comprises a plurality of internal steps (or steps) 220 which are regularly distributed around the longitudinal axis X-X. Each of these internal steps 220 extends, firstly longitudinally between the two longitudinal ends (upstream and downstream) of the inner wall, and secondly radially outwardly thereof.

En d'autres termes, la surface intérieure de la paroi interne 212 est profilée avec une pluralité de marches 220 faisant saillie vers l'extérieur de la paroi. Par ailleurs, on désigne par cavité interne 222 l'espacement circonférentiel qui est défini entre deux marches internes 220 adjacentes.In other words, the inner surface of the inner wall 212 is profiled with a plurality of steps 220 protruding outwardly from the wall. Furthermore, internal cavity 222 designates the circumferential spacing that is defined between two adjacent internal steps 220.

De même, la paroi externe 214 de la chambre de combustion comporte une pluralité de marches (ou marches) externes 224 régulièrement réparties autour de l'axe longitudinal X-X. Chaque marche externe 224 s'étend, d'une part longitudinalement entre les deux extrémités longitudinales de la paroi externe, et d'autre part radialement vers l'intérieur de celle-ci.Similarly, the outer wall 214 of the combustion chamber comprises a plurality of steps (or steps) external 224 evenly distributed around the longitudinal axis X-X. Each external step 224 extends, firstly longitudinally between the two longitudinal ends of the outer wall, and secondly radially inwardly thereof.

De manière analogue à la paroi interne, la surface extérieure de la paroi externe 214 est profilée avec une pluralité de marches 224 faisant saillie vers l'intérieur de la paroi. On désigne par cavité externe 226 l'espacement circonférentiel qui est défini entre deux marches externes 224 adjacentes.In a similar manner to the inner wall, the outer surface of the outer wall 214 is profiled with a plurality of steps 224 projecting inwardly from the wall. External cavity 226 denotes the circumferential spacing that is defined between two adjacent external steps 224.

Toujours selon l'invention, certaines des cavités internes 222 et certaines des cavités externes 226 sont alimentées en carburant selon une direction sensiblement radiale.Still according to the invention, some of the internal cavities 222 and some of the external cavities 226 are supplied with fuel in a substantially radial direction.

A cet effet, la chambre de combustion 202 selon l'invention comprend encore une pluralité de systèmes d'injection de carburant 228 répartis sur les parois interne 212 et 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.For this purpose, the combustion chamber 202 according to the invention also comprises a plurality of fuel injection systems 228 distributed on the inner walls 212 and outer 214 around the longitudinal axis XX of the turbomachine and opening into the combustion chamber. combustion 216 in a substantially radial direction.

De façon plus précise, comme représenté sur les figures 2 et 3, les systèmes d'injection de carburant 228 débouchent radialement dans certaines des cavités internes 222 et certaines des cavité externes 226.More specifically, as shown on the figures 2 and 3 the fuel injection systems 228 radially open into some of the internal cavities 222 and some of the outer cavities 226.

Ainsi, dans l'exemple de réalisation des figures 2 à 4, les systèmes d'injection de carburant 228 débouchent dans toutes les cavités externes 226 et dans seulement une cavité interne 222 sur deux. Bien entendu, d'autres configurations sont possibles : toutes les cavités internes et toutes les cavités externes peuvent être alimentées en carburant ; seulement une cavité externe sur deux et toutes les cavités internes sont alimentées en carburant ; etc. Le principe régissant le choix de la configuration d'alimentation de ces cavités est de parvenir à optimiser les performances de la chambre de combustion pour chaque point du domaine de vol.Thus, in the exemplary embodiment of Figures 2 to 4 , the fuel injection systems 228 open into all the external cavities 226 and into only one internal cavity 222 out of two. Good Of course, other configurations are possible: all the internal cavities and all the external cavities can be supplied with fuel; only one external cavity out of two and all internal cavities are fueled; etc. The principle governing the choice of the supply configuration of these cavities is to optimize the performance of the combustion chamber for each point of the flight envelope.

De façon avantageuse, les systèmes d'injection de carburant 228 comportent des injecteurs pilote 228a alternant circonférentiellement avec des injecteurs plein gaz 228b.Advantageously, fuel injection systems 228 comprise pilot injectors 228a alternating circumferentially with full-throttle injectors 228b.

Ainsi, toujours dans l'exemple de réalisation des figures 2 à 4, les systèmes d'injection de carburant 228 alimentant les cavités externes 226 comportent bien une alternance d'injecteurs pilote 228a avec des injecteurs plein gaz, et les systèmes d'injection de carburant 228 alimentant les cavités internes 222 comportent des injecteurs plein gaz et des injecteurs pilote.So, again in the example of realization of Figures 2 to 4 , the fuel injection systems 228 supplying the external cavities 226 do indeed comprise an alternation of pilot injectors 228a with full-throttle injectors, and the fuel injection systems 228 supplying the internal cavities 222 comprise full-throttle injectors and injectors. pilot injectors.

Classiquement, les injecteurs pilote 228a assurent l'allumage et les phases de ralenti de la turbomachine et les injecteurs plein gaz 228b 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 de décollage ne sont alimentés qu'au-delà d'un certain régime déterminé.Conventionally, the pilot injectors 228a provide ignition and idle phases of the turbomachine and the full-throttle injectors 228b are involved in the take-off, climb and cruise phases. In general, the pilot injectors are fueled continuously while the takeoff injectors are only fed beyond a certain determined regime.

Selon une caractéristique particulière avantageuse de l'invention, les systèmes d'injection de carburant 228 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 228 do not have 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 228a sont du même type que les injecteurs plein gaz 228b.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 228a are of the same type as the full-throttle injectors 228b.

Par ailleurs, contrairement à l'exemple de réalisation des figures 2 à 4, les injecteurs plein gaz 228b peuvent être décalés axialement vers l'aval par rapport aux injecteurs pilote 228a.Moreover, contrary to the example of realization of Figures 2 to 4 , the full-throttle injectors 228b can be axially offset downstream relative to the pilot injectors 228a.

Toujours selon l'invention, au moins certaines des cavités internes 222 et certaines des cavités externes 226 sont alimentées en air extérieur à la chambre de combustion 202 selon une même direction sensiblement circonférentielle.Still according to the invention, at least some of the internal cavities 222 and some of the external cavities 226 are supplied with air outside the combustion chamber 202 in a same substantially circumferential direction.

A cet effet, les cavités internes 222 et externes 226 qui sont alimentées en air au moyen d'une pluralité d'orifices d'injection d'air 230 pratiquées dans une paroi sensiblement radiale 232 des marches internes 220 et externes 224 correspondantes. Ces orifices d'injection d'air 230 s'ouvrent vers l'extérieur de la chambre de combustion 202 et débouchent dans la cavité interne ou externe correspondante selon une direction sensiblement circonférentielle.For this purpose, the internal cavities 222 and external 226 which are supplied with air by means of a plurality of air injection orifices 230 formed in a substantially radial wall 232 of the internal 220 and external steps 224 corresponding. These air injection orifices 230 open towards the outside of the combustion chamber 202 and open into the corresponding internal or external cavity in a substantially circumferential direction.

Ainsi, dans l'exemple de réalisation des figures 2 à 4, toutes les cavités internes 222 et toutes les cavités externes 226 sont alimentées en air au moyen de tels orifices d'injection d'air (c'est-à-dire même les cavités internes qui ne sont pas alimentées en carburant). Bien entendu, d'autres configurations sont possibles en fonction des besoins : seules certaines des cavités internes et certaines des cavités externes peuvent être alimentées en air.Thus, in the exemplary embodiment of Figures 2 to 4 all internal cavities 222 and all external cavities 226 are supplied with air by means of such air injection ports (i.e. even internal cavities which are not fueled). Of course, other configurations are possible depending on the needs: only some of the internal cavities and some of the external cavities can be supplied with air.

Il est à noter que l'injection circonférentielle d'air dans le foyer de combustion 216 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 internes 222 et externes 226 de la chambre de combustion. Par ailleurs, le sens de rotation pour l'injection circonférentielle d'air dans ces cavités 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 into the combustion chamber 216 is made in the same direction of rotation (that of the needles of a watch for the embodiment of figures 2 and 3 ) for all internal cavities 222 and external 226 of the combustion chamber. Moreover, the direction of rotation for the circumferential injection of air into these cavities is that of the compressed air coming from the compression section of the turbomachine.

Il est encore à noter que l'alimentation en air du foyer de combustion 206 est uniquement réalisée au moyen des orifices d'injection d'air 230 débouchant dans certaines des cavités internes et externes selon une direction circonférentielle (une très faible part d'air pénètre également dans le foyer de combustion en passant par des trous de multiperforation pratiquées dans les parois 212, 214 et 218 de la chambre de combustion pour le refroidissement de ces parois, ces trous n'étant pas représentés sur les figures).It should also be noted that the air supply of the combustion chamber 206 is only achieved by means of the air injection orifices 230 opening into some of the internal and external cavities in a circumferential direction (a very small proportion of air also enters the combustion chamber through multiperforation holes made in the walls 212, 214 and 218 of the combustion chamber for cooling these walls, these holes are not shown in the figures).

Enfin, les cavités internes et externes qui sont alimentées en carburant ne sont pas forcément homogènes en ce qui concerne leur dimension radiale (c'est-à-dire la hauteur de la marche correspondante) et circonférentielle de manière à pouvoir faire varier le temps de résidence suivant la cavité considérée. De même, comme représenté sur la figure 4, la hauteur des marches n'est pas forcément constante sur toute la longueur de la paroi (c'est-à-dire entre ses extrémités amont et aval). En outre, le débit d'air alimentant ces cavités peut varier suivant la cavité considérée.Finally, the internal and external cavities that are supplied with fuel are not necessarily homogeneous with respect to their radial dimension (that is to say the height of the corresponding step) and circumferential so as to be able to vary the residence time following the cavity considered. Similarly, as shown on the figure 4 , the height of the steps is not necessarily constant over the entire length of the wall (that is to say between its upstream and downstream ends). In addition, the air flow supplying these cavities may vary depending on the cavity considered.

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 contourne la chambre de combustion 202 et alimente certaines au moins des cavités internes 222 et externes 226 après avoir refroidi les parois et enveloppes de la chambre de combustion. Cet air est injecté dans ces cavités par l'intermédiaire des orifices d'injection d'air 230 selon le sens de rotation de l'air à son entrée dans la section de combustion. Dans certaines de ces cavités alimentées en air, l'air est mélangé et brûlé au carburant injecté par les systèmes d'injection de carburant 228.The operation of the combustion chamber is as follows: the compressed air coming from the compression section 100 and rotating about the longitudinal axis XX enters the combustion section 200. This air bypasses the combustion chamber 202 and supplies at least some of the internal and external cavities 222 and 226 after cooling the walls and casings of the combustion chamber. This air is injected into these cavities via the air injection orifices 230 in the direction of rotation of the air at its entry into the combustion section. In some of these air-fed cavities, the air is mixed and burned with the fuel injected by the fuel injection systems 228.

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, les marches internes 220 et externes 224 de la chambre de combustion comportent une autre paroi 232' (opposée à celle 232 munie d'orifices d'injection d'air) qui s'étend selon une direction sensiblement circonférentielle et qui présente, en coupe transversale, une section sensiblement curviligne (contrairement à la paroi 232 qui est sensiblement plane et radiale). La courbure de cette paroi permet de former une rampe pour accompagner le mouvement de rotation de l'air injecté dans les cavités par les orifices d'injection d'air 230. Bien entendu, toute autre forme de paroi (plane ou curviligne) est envisageable.In the exemplary embodiment of figures 2 and 3 the internal and external steps 224 224 of the combustion chamber comprise another wall 232 '(opposite to that 232 provided with air injection orifices) which extends in a substantially circumferential direction and which has, in section transverse, a substantially curvilinear section (unlike the wall 232 which is substantially flat and radial). The curvature of this wall makes it possible to form a ramp to accompany the rotational movement of the air injected into the cavities via the air injection orifices 230. Of course, any other wall form (planar or curvilinear) can be envisaged. .

De manière générale, le nombre et les dimensions géométriques des cavités internes et externes 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 dans ces cavités, ainsi que de la position circonférentielle relative des systèmes d'injection de carburant par rapport aux marches internes et externes.In general, the number and the geometrical dimensions of the internal and external cavities 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 in these cavities, as well as the relative circumferential position of the fuel injection systems with respect to the internal and external steps.

Enfin, comme représenté sur les figures 1 à 4, la paroi interne 212 et la paroi externe 214 de la chambre de combustion peuvent chacune comporter à leur extrémité aval une bride annulaire, respectivement 234 et 236, qui est munie d'une pluralité de trous 238 régulièrement répartis autour de l'axe longitudinal X-X et destinés à alimenter en air de refroidissement la section de turbine 300.Finally, as represented on Figures 1 to 4 , the inner wall 212 and the outer wall 214 of the combustion chamber may each comprise at their downstream end an annular flange, respectively 234 and 236, which is provided with a plurality of holes 238 regularly distributed about the longitudinal axis XX and for supplying cooling air to the turbine section 300.

Claims (7)

  1. A turbomachine combustion chamber (202) comprising:
    . an inner annular wall (212) of longitudinal axis (X-X);
    . an outer annular wall (214) centered on the longitudinal axis and surrounding the inner wall so as to co-operate therewith to define an annular space (216) forming a combustion area;
    . a transverse annular wall (218) transversely interconnecting the upstream longitudinal ends of the inner and outer walls; and
    . a plurality of fuel injection systems (228);
    . the inner wall (212) including a plurality of inner steps (220) that are regularly distributed around the longitudinal axis, each inner step extending longitudinally between the two longitudinal ends of the inner wall and radially towards the outside thereof, the circumferential spacing between two adjacent inner steps defining an inner cavity (222);
    · the outer wall (214) including a plurality of outer steps (224) that are regularly distributed around the longitudinal axis, each outer step extending longitudinally between the two longitudinal ends of the outer wall and radially towards the inside thereof, the circumferential spacing between two adjacent outer steps defining an outer cavity (226); and
    · at least some of the inner and outer cavities are fed with air external to the combustion chamber in a common direction that is substantially circumferential, characterized in that these inner and outer cavities are also fed with fuel in a direction that is substantially radial.
  2. A combustion chamber according to claim 1, in which some of the inner and outer steps (220 and 224) include respective substantially radial walls (232), each provided with a plurality of air injection orifices (230) opening to the outside of the combustion chamber and into the adjacent inner or outer cavity.
  3. A combustion chamber according to claim 2, in which each of the inner and outer steps (220 and 224) includes a respective other wall (232') that presents, in cross-section, a section that is substantially curvilinear.
  4. A combustion chamber according to any one of claims 1 to 3, in which the fuel injection systems (228) comprise pilot injectors (228a) alternating circumferentially with full-throttle injectors (228b).
  5. A combustion chamber according to claim 4, in which the full-throttle injectors (228b) are offset axially downstream relative to the pilot injectors (228a).
  6. A combustion chamber according to any one of claims 1 to 5, in which the fuel injection systems (228) do not include associated air systems.
  7. A turbomachine characterized in that includes a combustion chamber (202) according to any one of claims 1 to 6.
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FR0757356A FR2920523B1 (en) 2007-09-05 2007-09-05 TURBOMACHINE COMBUSTION CHAMBER WITH AIR HELICOIDAL CIRCULATION.

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FR2920523A1 (en) 2009-03-06
EP2034245A1 (en) 2009-03-11
US7614234B2 (en) 2009-11-10
CA2639356A1 (en) 2009-03-05
US20090056338A1 (en) 2009-03-05
RU2008135874A (en) 2010-03-10
JP5214375B2 (en) 2013-06-19
FR2920523B1 (en) 2009-12-18
RU2484377C2 (en) 2013-06-10
CN101382297A (en) 2009-03-11
JP2009063287A (en) 2009-03-26
DE602008001042D1 (en) 2010-06-02
CN101382297B (en) 2011-11-23
CA2639356C (en) 2015-06-23

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