EP3247945A1 - Fuel injection system for aircraft turbomachine, comprising a variable section air through duct - Google Patents

Fuel injection system for aircraft turbomachine, comprising a variable section air through duct

Info

Publication number
EP3247945A1
EP3247945A1 EP16703587.2A EP16703587A EP3247945A1 EP 3247945 A1 EP3247945 A1 EP 3247945A1 EP 16703587 A EP16703587 A EP 16703587A EP 3247945 A1 EP3247945 A1 EP 3247945A1
Authority
EP
European Patent Office
Prior art keywords
injection system
central body
flared
air
combustion chamber
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
EP16703587.2A
Other languages
German (de)
French (fr)
Other versions
EP3247945B1 (en
Inventor
Yoann Mery
Olivier BIDART
Julien LEPAROUX
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
Safran Aircraft Engines SAS
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Safran Aircraft Engines SAS filed Critical Safran Aircraft Engines SAS
Publication of EP3247945A1 publication Critical patent/EP3247945A1/en
Application granted granted Critical
Publication of EP3247945B1 publication Critical patent/EP3247945B1/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23N—REGULATING OR CONTROLLING COMBUSTION
    • F23N1/00—Regulating fuel supply
    • F23N1/02—Regulating fuel supply conjointly with air supply
    • F23N1/025—Regulating fuel supply conjointly with air supply using electrical or electromechanical means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23N—REGULATING OR CONTROLLING COMBUSTION
    • F23N1/00—Regulating fuel supply
    • F23N1/02—Regulating fuel supply conjointly with air supply
    • F23N1/027—Regulating fuel supply conjointly with air supply using mechanical means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/30—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D2900/00—Special features of, or arrangements for burners using fluid fuels or solid fuels suspended in a carrier gas
    • F23D2900/11101—Pulverising gas flow impinging on fuel from pre-filming surface, e.g. lip atomizers

Definitions

  • the present invention relates to the field of combustion chambers for aircraft turbomachines, preferably for turbojet engines.
  • Injection systems are the subject of many developments. Their design is constantly optimized to improve their performance in ground ignition, reignition at altitude, or extinction. It is also desirable to limit idle pollution as much as possible, which is closely related to the capacity of the injection system to be atomized and to mix the injected fuel with the air.
  • Such aerodynamic injection systems are known from documents FR 2 875 585, or from documents FR 2 685 452 and FR 2 832 493.
  • the invention thus aims to at least partially overcome the disadvantages relating to the embodiments of the prior art.
  • the subject of the invention is an assembly comprising an injection system for an aircraft turbomachine combustion chamber, as well as a fuel injector cooperating with the injection system, this assembly being in accordance with the claim 1.
  • the invention is first of all remarkable in that it makes it possible to vary the length of the fuel film that matches the central body of the injection system, as a function of the output range of this injection system in the aerodynamic bowl.
  • This ability to vary the length of the fuel film advantageously influences the stability of the flame, which can be satisfactory for all engine speeds.
  • the invention introduces an additional degree of freedom in the design of the injection system, by varying the passage section of the air passage channel defined between the first and second coaxial flared ends. Thanks to this feature, the geometry of the injection system can be adapted according to the operating points of the turbomachine, which also contributes to obtaining increased performance for all engine speeds.
  • the fact of being able to vary the passage section of the air passage channel makes it possible to influence the richness of the air-fuel mixture, which directly impacts the stability of this mixture.
  • this faculty of variation of the passage section of the air passage channel can influence the atomization of fuel, which occurs at the output of the second flared end of the central body.
  • the atomization phenomenon has a direct impact on the stability of the combustion chamber, on the ground ignition and reignition capability at altitude, or on pollutant emissions at idle speed. All of these parameters can thus be optimized for all operating points of the turbomachine, thanks to the degree of freedom of movement introduced into the design of the injection system according to the invention.
  • the invention preferably has at least one of the following optional features, taken singly or in combination.
  • first and second flared ends are of frustoconical shape and delimit between them a frustoconical air passage channel of variable section as a function of an axial relative position between said first and second flared ends. Nevertheless, other flared non-frustoconical shapes can be retained, without departing from the scope of the invention.
  • the injection system comprises an intermediate structure arranged radially between a base of the central body and a base of the aerodynamic bolus, said intermediate structure delimiting with the base of the central body an axial flow channel of the fuel film towards said second end flared from the central body.
  • the base of the aerodynamic bowl comprises two concentric walls between which is arranged an air introduction auger between the two concentric walls.
  • Said moving means comprise a motor, for example a linear motor.
  • Said first flared end of the aerodynamic bowl is pierced with air introduction holes in a combustion chamber delimited by this bowl.
  • the subject of the invention is also an aircraft turbomachine combustion chamber comprising a chamber bottom pierced with apertures spaced apart from one another, the combustion chamber comprising, associated with each opening of the chamber bottom, a unit such that described above.
  • the subject of the invention is an aircraft turbomachine comprising such a combustion chamber.
  • FIG. 1 shows a schematic longitudinal sectional view of a turbojet according to the invention
  • FIG. 2 shows a longitudinal half-sectional view of the combustion chamber of the turbojet shown in the previous figure
  • FIG. 3a and 3b show views of the principle of an injection system equipping the combustion chamber shown in the previous figure, respectively in two separate positions of the central body of the injection system;
  • FIGS. 4a and 4b show views of an injection system according to a preferred embodiment of the invention, respectively in two distinct positions of the central body of this injection system;
  • FIG. 5 shows a similar to those of Figures 4a and 4b, in which it was partially schematized the air and fuel paths through the injection system.
  • an aircraft turbomachine 1 according to a preferred embodiment of the invention.
  • This is a turbojet engine with double flow and double body.
  • turbomachine of another type for example a turboprop, without departing from the scope of the invention.
  • the turbomachine 1 has a longitudinal axis 3 around which its various components extend. It comprises, from upstream to downstream in a main direction of gas flow through this turbomachine, a fan 2, a low pressure compressor 4, a high pressure compressor 6, a combustion chamber 8, a high pressure turbine 10 and a low-pressure turbine 12.
  • this turbomachine 1 is controlled by a control unit 13, only shown schematically. This unit 13 allows in particular to control the different operating points of the turbomachine.
  • Part of the combustion chamber 8 is reproduced in more detail in Figure 2. It has in particular an outer shell 14 centered on the axis 3, an inner ring 16 also centered on the same axis, and a chamber bottom 18 connecting the two ferrules at their upstream end.
  • Fuel injectors 20 are regularly distributed over the chamber bottom, in the circumferential direction (a single injector being visible in FIG. 2). Each of them has an injector nose 21, oriented along a main axis 22 slightly inclined with respect to the axis 3. In this respect, it is indicated that this axis 22 is parallel to the main flow direction of the flow 24 through the room.
  • Each injector 20 is associated with an injection system 30, shown diagrammatically in FIG. 2.
  • the injection system 30 cooperates upstream with the injector nozzle 21, while it opens downstream into the combustion chamber 8
  • the injection system 30 is housed in an opening 32 formed through the chamber bottom 18.
  • FIGS 3a and 3b show the principle of the injection system 30 according to the invention.
  • This system 30, of the aerodynamic injection system type first comprises an outer wall formed by an aerodynamic bolus 40, equipped with a first end flared downstream 42, said divergent portion.
  • This flared end 42 is of frustoconical shape, of axis 22.
  • the bowl comprises a base 44 also centered on the axis 22.
  • the system 30 comprises a central body 46 solid at least partially housed in the housing. The interior of the space defined by the bowl 40.
  • the body 46 is equipped with a second end flared downstream 48, said divergent portion.
  • This flared end 48 is of frustoconical shape, of axis 22.
  • the body comprises a base 50 also centered on the axis 22, and arranged inside the base 44 mentioned above.
  • the injector 20 cooperates with the injection system 30 so that a film of fuel travels along the central body 46, downstream.
  • the fuel film 52 thus flows downstream on the outer surface of the base 50 and the flared end 48 of the central body 46.
  • the film 52 is atomized. which allows it to hang the flame 54 located inside the chamber.
  • the recirculation formed at the end of this diverging portion 48 stabilizes the flame and thus increase the extinction performance of the hearth.
  • the flared end 42 has an annular row of holes 66 for introducing air into the combustion chamber. These holes are located near a fixing flange (not shown) for fixing the bowl on the chamber floor 18, in the associated opening 32.
  • the selected design thus implements a circulation of a fuel film 52 along the central body 46 of the injection system, as is for example the prior art.
  • This fuel film injection design differs from the so-called "spray" design in which the fuel is injected via a spin, also allowing the passage of air. Due to the passage of fuel in the spin, in the form of a spray, the air permeability of the injection system is modified. In contrast, in the invention, the air is intended to circulate through the injection system 30 via an air passage channel 56 delimited between the bowl 40 and the central body 46. This circulation, preferably initiated by a twist (not shown in Figures 3a and 3b), is not disturbed by the fuel film 48 flowing only on the inner wall of the channel 56.
  • a frustoconical air passage channel 60 constituting the downstream portion of said channel 56.
  • the frustoconical channel 60 is centered on the axis 22 and has a cross section referenced SI in Figure 3a.
  • One of the peculiarities of the invention resides in the fact that the injection system integrates a degree of freedom of movement making it possible to vary the cross section of the frustoconical channel 60, according to the needs met.
  • the injection system 30 comprises moving means 62, allowing relative movement between the first and second flared ends 42, 48, along the axis 22.
  • These means 62 are of type conventional, for example incorporating a linear motor, or an electromagnet. They are controlled by the unit 13, and make it possible to set the central body 46 in motion inside the bowl 40, the latter being fixed with respect to the chamber bottom 18 and to the injector 20. Also, depending on the axial relative position between the first and second flared ends 42, 48, the cross section of the channel 60 varies. In FIG. 3b, this section referenced S2 is smaller than the section S1 of FIG. 3a, because the central body 46 has been moved upstream by the means 62.
  • the fact of being able to vary the passage section of the channel 60 makes it possible to influence the richness of the air-fuel mixture, which directly impacts the stability of this mixture.
  • This faculty of variation of the passage section of the air passage channel makes it possible to influence the atomization of the fuel, which occurs at the outlet of the second flared end of the central body.
  • the atomization can be characterized by the ratio of the amounts of air and fuel movements, and therefore directly dependent on the passage section of the air passage channel.
  • This atomization may also vary according to the length of the fuel film 52 conforming externally to the central body 46, this length being greater in the position of FIG. 3a than in the position of FIG. 3b on which the central body 46 is in position. withdrawal, upstream.
  • the atomization phenomenon has a direct impact on the stability of the combustion chamber, on the ground ignition and reignition capability at altitude, or on pollutant emissions at idle speed. All of these parameters can thus be optimized for all operating points of the turbomachine. For example, a significant pressure drop of air through the injection system is advantageous for atomizing and mixing the fuel at idle, in order to increase the stability of the flame.
  • the position of FIG. 3b, with the reduced section S2 will therefore be preferred for this idling speed of the turbomachine.
  • the position of FIG. 3a will preferably be retained for the full throttle and cruising regimes.
  • FIG. 4a and 4b there is shown the injection system 30 according to a preferred embodiment of the invention.
  • the elements bearing the same reference numbers as elements of the principle figures 3a and 3b correspond to identical or similar elements.
  • the moving means 62 of the main body 46 have not been shown in these figures 4a and 4b. Nevertheless, these means 62 are obviously provided and controlled to move the main body 46 from the position of Figure 4a to that of Figure 4b, and vice versa.
  • the injection system 30 comprises an intermediate structure 70, arranged radially between the base 50 of the central body 46 and the base 44 of the bowl 40. It is relatively to this intermediate structure 70 that the main body 46 is capable of being displaced axially between the two positions of FIGS. 4a and 4b, while the structure 70 remains fixed relative to the injector 20 and the bowl 40.
  • the intermediate structure 70 defines with the outer surface of the base 50 an axial annular channel 72 for the flow of the fuel film 52, in the direction of the second flared end 48 of the central body 46. It is indeed this channel 72 which is fed in known manner by the injector 20 and which allows to generate the thin film of fuel 52 along the central body 46, before encountering the air introduced into the injection system. In the preferred embodiment shown, the fuel bypasses the upstream end of the solid central body 46 before marrying the outer wall thereof internally defining the axial annular channel 72.
  • the base 44 of the bowl 40 here comprises two concentric walls 44a, 44b between which is arranged a swirl 76 for introducing air between the two concentric walls, this swirl being of axial or radial character.
  • internal wall 44b surrounds the intermediate structure 70, so as to delimit between them a channel 80 opening downstream.
  • the channel 80 is not intended to be traversed by an air flow.
  • the latter In the channel 60 of greater width than that of the channel 72 in which the fuel film 52 is created, the latter remains confined along the lateral surface of the frustoconical end 48 of the central body 46, thanks to the passage of the air 82 in the same channel 60.
  • the darkest shaded portion 52 of Figure 5 represents the fuel path from the injector 20 to the flame 54, while the lightest gray portion 82 represents the flow of the fuel. air.

Abstract

The invention relates to an assembly comprising an injection system (30) and an injector (20) for an aircraft turbomachine combustion chamber (8), the system (30) comprising an aerodynamic bowl (40) comprising a first end widening toward the downstream end (42) and centred on a central axis (22) of the injection system, this also comprising a central body (46) along which a film of fuel (52) is intended to flow in the downstream direction. According to the invention, the central body (46) comprises a second end widening toward the downstream end (48), the first and second widening ends (42, 48) between them delimiting an air through duct (60) and the system (30) comprises motion inducing means (62) allowing a relative movement between the first widening end (42) which is stationary and the second widening end (48), along the central axis (22) of the injection system, by moving the central body (46) relative to the injector (20).

Description

SYSTEME D'INJECTION DE CARBURANT POUR TURBOMACHINE D'AERONEF, COMPRENANT UN CANAL DE TRAVERSEE D'AIR A SECTION VARIABLE  FUEL INJECTION SYSTEM FOR AIRCRAFT TURBINE ENGINE COMPRISING A VARIABLE SECTION AIR AIR CHANNEL
DESCRIPTION DESCRIPTION
DOMAINE TECHNIQUE TECHNICAL AREA
La présente invention se rapporte au domaine des chambres de combustion pour turbomachines d'aéronef, de préférence pour des turboréacteurs. The present invention relates to the field of combustion chambers for aircraft turbomachines, preferably for turbojet engines.
Elle se rapporte plus précisément aux systèmes d'injection équipant la chambre de combustion, ces systèmes d'injection ayant pour fonction principale de mélanger l'air au carburant délivré par les injecteurs.  It relates more precisely to the injection systems fitted to the combustion chamber, these injection systems having the main function of mixing the air fuel delivered by the injectors.
ÉTAT DE LA TECHNIQUE ANTÉRIEURE STATE OF THE PRIOR ART
Les systèmes d'injection font l'objet de nombreux développements. Leur conception est sans cesse optimisée de façon à améliorer leurs performances en allumage au sol, en rallumage en altitude, ou encore en extinction. I l est également recherché à limiter autant que possible la pollution au ralenti, qui est étroitement liée à la capacité du système d'injection à atomiser et mélanger le carburant injecté avec l'air. De tels systèmes d'injection aérodynamiques sont connus des documents FR 2 875 585, ou encore des documents FR 2 685 452 et FR 2 832 493. Injection systems are the subject of many developments. Their design is constantly optimized to improve their performance in ground ignition, reignition at altitude, or extinction. It is also desirable to limit idle pollution as much as possible, which is closely related to the capacity of the injection system to be atomized and to mix the injected fuel with the air. Such aerodynamic injection systems are known from documents FR 2 875 585, or from documents FR 2 685 452 and FR 2 832 493.
Néanmoins, l'obtention de performances optimales pour certains points de fonctionnement de la turbomachine, passe par des contraintes de dimensionnement qui peuvent se révéler inadaptées pour d'autres points de fonctionnement, da ns lesquels les performances globales sont alors réduites. En effet, les contraintes de dimensionnement pour le fonctionnement au plein gaz et en croisière peuvent fortement différer de celles du fonctionnement à bas régime. Par exemple, une perte de charge importante de l'air à travers le système d'injection est avantageuse pour atomiser et mélanger le carburant au ralenti, afin d'augmenter la stabilité de la flamme. Cependant, cette perte de charge devient pénalisante en termes de consommation spécifique en fonctionnement au plein gaz et en croisière. However, obtaining optimum performance for certain points of operation of the turbomachine, requires sizing constraints that may be unsuitable for other operating points, in which the overall performance is reduced. Indeed, the design constraints for full throttle operation and cruising may differ significantly from those of low speed operation. For example, a significant pressure drop of air through the injection system is advantageous for atomizing and mixing the fuel at idle, in order to increase the stability of the flame. However, this pressure drop becomes penalizing in terms of specific consumption in full throttle operation and cruising.
EXPOSÉ DE L'INVENTION STATEMENT OF THE INVENTION
L'invention a ainsi pour but de remédier au moins partiellement aux inconvénients relatifs aux réalisations de l'art antérieur. The invention thus aims to at least partially overcome the disadvantages relating to the embodiments of the prior art.
Pour ce faire, l'invention a pour objet un ensemble comprenant un système d'injection pour chambre de combustion de turbomachine d'aéronef, ainsi qu'un injecteur de carburant coopérant avec le système d'injection, cet ensemble étant conforme à la revendication 1.  To do this, the subject of the invention is an assembly comprising an injection system for an aircraft turbomachine combustion chamber, as well as a fuel injector cooperating with the injection system, this assembly being in accordance with the claim 1.
L'invention est tout d'abord remarquable en ce qu'elle permet de faire varier la longueur du film de carburant qui épouse le corps central du système d'injection, en fonction de l'étendue de sortie de ce système d'injection dans le bol aérodynamique. Cette capacité à faire varier la longueur du film de carburant influe avantageusement sur la stabilité de la flamme, qui peut ainsi être satisfaisante pour tous les régimes moteur.  The invention is first of all remarkable in that it makes it possible to vary the length of the fuel film that matches the central body of the injection system, as a function of the output range of this injection system in the aerodynamic bowl. This ability to vary the length of the fuel film advantageously influences the stability of the flame, which can be satisfactory for all engine speeds.
De plus, l'invention introduit un degré de liberté supplémentaire da ns la conception du système d'injection, en permettant de faire varier la section de passage du canal de traversée d'air défini entre les première et seconde extrémités évasées coaxiales. Grâce à cette particularité, la géométrie du système d'injection peut être adaptée en fonction des points de fonctionnement de la turbomachine, ce qui contribue également à l'obtention de performances accrues pour l'ensemble des régimes moteur. En particulier, le fait de pouvoir faire varier la section de passage du canal de traversée d'air permet d'influencer la richesse du mélange air-carburant, qui impacte directement la stabilité de ce mélange. En outre, cette faculté de variation de la section de passage du canal de traversée d'air permet d'influencer l'atomisation du carburant, qui se produit en sortie de la seconde extrémité évasée du corps central. Avantageusement, le phénomène d'atomisation présente un impact direct sur la stabilité de la chambre de combustion, sur la capacité d'allumage au sol et de rallumage en altitude, ou encore sur les émissions polluantes en régime de ralenti. L'ensemble de ces paramètres peuvent ainsi être optimisés pour tous les points de fonctionnement de la turbomachine, grâce au degré de liberté de mouvement introduit dans la conception du système d'injection selon l'invention. In addition, the invention introduces an additional degree of freedom in the design of the injection system, by varying the passage section of the air passage channel defined between the first and second coaxial flared ends. Thanks to this feature, the geometry of the injection system can be adapted according to the operating points of the turbomachine, which also contributes to obtaining increased performance for all engine speeds. In particular, the fact of being able to vary the passage section of the air passage channel makes it possible to influence the richness of the air-fuel mixture, which directly impacts the stability of this mixture. In addition, this faculty of variation of the passage section of the air passage channel can influence the atomization of fuel, which occurs at the output of the second flared end of the central body. Advantageously, the atomization phenomenon has a direct impact on the stability of the combustion chamber, on the ground ignition and reignition capability at altitude, or on pollutant emissions at idle speed. All of these parameters can thus be optimized for all operating points of the turbomachine, thanks to the degree of freedom of movement introduced into the design of the injection system according to the invention.
L'invention présente de préférence au moins l'une des caractéristiques optionnelles suivantes, prises isolément ou en combinaison.  The invention preferably has at least one of the following optional features, taken singly or in combination.
Lesdites première et seconde extrémités évasées sont de forme tronconique et délimitent entre elles un canal tronconique de traversée d'air, de section variable en fonction d'une position relative axiale entre lesdites première et seconde extrémités évasées. Néanmoins, d'autres formes évasées non tronconiques peuvent être retenues, sans sortir du cadre de l'invention.  Said first and second flared ends are of frustoconical shape and delimit between them a frustoconical air passage channel of variable section as a function of an axial relative position between said first and second flared ends. Nevertheless, other flared non-frustoconical shapes can be retained, without departing from the scope of the invention.
Le système d'injection comporte une structure intermédiaire agencée radialement entre une base du corps central et une base du bol aérodynamique, ladite structure intermédiaire délimitant avec la base du corps central un canal axial d'écoulement du film de carburant en direction de ladite seconde extrémité évasée du corps central.  The injection system comprises an intermediate structure arranged radially between a base of the central body and a base of the aerodynamic bolus, said intermediate structure delimiting with the base of the central body an axial flow channel of the fuel film towards said second end flared from the central body.
La base du bol aérodynamique comporte deux parois concentriques entre lesquelles est agencée une vrille d'introduction d'air entre les deux parois concentriques.  The base of the aerodynamic bowl comprises two concentric walls between which is arranged an air introduction auger between the two concentric walls.
Lesdits moyens de mise en mouvement comprennent un moteur, par exemple un moteur linéaire.  Said moving means comprise a motor, for example a linear motor.
Ladite première extrémité évasée du bol aérodynamique est percée de trous d'introduction d'air dans un foyer de combustion délimité par ce bol.  Said first flared end of the aerodynamic bowl is pierced with air introduction holes in a combustion chamber delimited by this bowl.
L'invention a également pour objet une chambre de combustion de turbomachine d'aéronef comprenant un fond de chambre percé d'ouvertures espacées les unes des autres, la chambre de combustion comprenant, associé à chaque ouverture du fond de chambre, un ensemble tel que décrit ci-dessus.  The subject of the invention is also an aircraft turbomachine combustion chamber comprising a chamber bottom pierced with apertures spaced apart from one another, the combustion chamber comprising, associated with each opening of the chamber bottom, a unit such that described above.
Enfin, l'invention a pour objet une turbomachine d'aéronef comprenant une telle chambre de combustion.  Finally, the subject of the invention is an aircraft turbomachine comprising such a combustion chamber.
D'autres avantages et caractéristiques de l'invention apparaîtront dans la description détaillée non limitative ci-dessous. BRÈVE DESCRIPTION DES DESSINS Other advantages and features of the invention will become apparent in the detailed non-limiting description below. BRIEF DESCRIPTION OF THE DRAWINGS
L'invention pourra être mieux comprise à la lecture de la description détaillée qui va suivre, d'exemples de mise en œuvre non limitatifs de celle-ci, ainsi qu'à l'examen des dessins annexés parmi lesquels ; The invention will be better understood on reading the following detailed description, non-limiting examples of implementation thereof, as well as on examining the appended drawings among which:
- la figure 1 représente une vue schématique en coupe longitudinale d'un turboréacteur selon l'invention ;  - Figure 1 shows a schematic longitudinal sectional view of a turbojet according to the invention;
- la figure 2 représente une vue en demi-coupe longitudinale de la chambre de combustion du turboréacteur montré sur la figure précédente ;  - Figure 2 shows a longitudinal half-sectional view of the combustion chamber of the turbojet shown in the previous figure;
- les figures 3a et 3b représentent des vues de principe d'un système d'injection équipant la chambre de combustion montrée sur la figure précédente, respectivement dans deux positions distinctes du corps central de ce système d'injection ;  - Figures 3a and 3b show views of the principle of an injection system equipping the combustion chamber shown in the previous figure, respectively in two separate positions of the central body of the injection system;
- les figures 4a et 4b représentent des vues d'un système d'injection selon un mode de réalisation préféré de l'invention, respectivement dans deux positions distinctes du corps central de ce système d'injection ; et  FIGS. 4a and 4b show views of an injection system according to a preferred embodiment of the invention, respectively in two distinct positions of the central body of this injection system; and
- la figure 5 représente une similaire à celles des figures 4a et 4b, sur laquelle il a été partiellement schématisé les trajets de l'air et du carburant à travers le système d'injection.  - Figure 5 shows a similar to those of Figures 4a and 4b, in which it was partially schematized the air and fuel paths through the injection system.
EXPOSÉ DÉTAILLÉ DE MODES DE RÉALISATION PARTICULIERS DETAILED PRESENTATION OF PARTICULAR EMBODIMENTS
En référence tout d'abord à la figure 1, il est représenté une turbomachine 1 d'aéronef, selon un mode de réalisation préféré de l'invention. Il s'agit ici d'un turboréacteur à double flux et à double corps. Néanmoins, il pourrait s'agir d'une turbomachine d'un autre type, par exemple un turbopropulseur, sans sortir du cadre de l'invention. Referring firstly to Figure 1, there is shown an aircraft turbomachine 1, according to a preferred embodiment of the invention. This is a turbojet engine with double flow and double body. Nevertheless, it could be a turbomachine of another type, for example a turboprop, without departing from the scope of the invention.
La turbomachine 1 présente un axe longitudinal 3 autour duquel s'étendent ses différents composants. Elle comprend, d'amont en aval selon une direction principale d'écoulement des gaz à travers cette turbomachine, une soufflante 2, un compresseur basse pression 4, un compresseur haute pression 6, une chambre de combustion 8, une turbine haute pression 10 et une turbine basse pression 12. De manière classique, cette turbomachine 1 est commandée par une unité de commande 13, uniquement représentée schématiquement. Cette unité 13 permet notamment de commander les différents points de fonctionnement de la turbomachine. The turbomachine 1 has a longitudinal axis 3 around which its various components extend. It comprises, from upstream to downstream in a main direction of gas flow through this turbomachine, a fan 2, a low pressure compressor 4, a high pressure compressor 6, a combustion chamber 8, a high pressure turbine 10 and a low-pressure turbine 12. In a conventional manner, this turbomachine 1 is controlled by a control unit 13, only shown schematically. This unit 13 allows in particular to control the different operating points of the turbomachine.
Une partie de la chambre de combustion 8 est reproduite de façon plus détaillée sur la figure 2. Elle présente en particulier une virole extérieure 14 centrée sur l'axe 3, une virole intérieure 16 également centrée sur ce même axe, et un fond de chambre 18 reliant les deux viroles à leur extrémité amont. Des injecteurs de carburant 20 sont régulièrement répartis sur le fond de chambre, selon la direction circonférentielle (un seul injecteur étant visible sur la figure 2). Chacun d'eux présente un nez d'injecteur 21, orienté selon un axe principal 22 légèrement incliné par rapport à l'axe 3. A cet égard, il est indiqué que cet axe 22 est parallèle à la direction principale d'écoulement du flux 24 à travers la chambre.  Part of the combustion chamber 8 is reproduced in more detail in Figure 2. It has in particular an outer shell 14 centered on the axis 3, an inner ring 16 also centered on the same axis, and a chamber bottom 18 connecting the two ferrules at their upstream end. Fuel injectors 20 are regularly distributed over the chamber bottom, in the circumferential direction (a single injector being visible in FIG. 2). Each of them has an injector nose 21, oriented along a main axis 22 slightly inclined with respect to the axis 3. In this respect, it is indicated that this axis 22 is parallel to the main flow direction of the flow 24 through the room.
A chaque injecteur 20 est associé un système d'injection 30, représenté schématiquement sur la figure 2. Le système d'injection 30 coopère en amont avec le nez d'injecteur 21, tandis qu'il débouche en aval dans la chambre de combustion 8. Le système d'injection 30 est logé dans une ouverture 32 pratiquée à travers le fond de chambre 18. Ainsi, sur cette chambre, il est prévu plusieurs ouvertures 32 espacées circonférentiellement les unes des autres, et chacune associée à un système d'injection 30 dont l'axe central correspond à l'axe 22.  Each injector 20 is associated with an injection system 30, shown diagrammatically in FIG. 2. The injection system 30 cooperates upstream with the injector nozzle 21, while it opens downstream into the combustion chamber 8 The injection system 30 is housed in an opening 32 formed through the chamber bottom 18. Thus, on this chamber, there are provided several openings 32 circumferentially spaced from each other, and each associated with an injection system 30 whose central axis corresponds to the axis 22.
Les figures 3a et 3b montrent le principe du système d'injection 30 selon l'invention. Ce système 30, du type système d'injection aérodynamique, comporte tout d'abord une paroi extérieure formée par un bol aérodynamique 40, équipé d'une première extrémité évasée vers l'aval 42, dite partie divergente. Cette extrémité évasée 42 est de forme tronconique, d'axe 22. En aval, le bol comporte une base 44 également centrée sur l'axe 22. En outre, le système 30 comprend un corps central 46 plein au moins logé en partie à l'intérieur de l'espace défini par le bol 40. Le corps 46 est équipé d'une seconde extrémité évasée vers l'aval 48, dite partie divergente. Cette extrémité évasée 48 est de forme tronconique, d'axe 22. En aval, le corps comporte une base 50 également centrée sur l'axe 22, et agencée à l'intérieur de la base 44 précitée.  Figures 3a and 3b show the principle of the injection system 30 according to the invention. This system 30, of the aerodynamic injection system type, first comprises an outer wall formed by an aerodynamic bolus 40, equipped with a first end flared downstream 42, said divergent portion. This flared end 42 is of frustoconical shape, of axis 22. Downstream, the bowl comprises a base 44 also centered on the axis 22. In addition, the system 30 comprises a central body 46 solid at least partially housed in the housing. The interior of the space defined by the bowl 40. The body 46 is equipped with a second end flared downstream 48, said divergent portion. This flared end 48 is of frustoconical shape, of axis 22. Downstream, the body comprises a base 50 also centered on the axis 22, and arranged inside the base 44 mentioned above.
De manière connue, l'injecteur 20 coopère avec le système d'injection 30 de façon à ce qu'un film de carburant chemine le long du corps central 46, vers l'aval. Le film de carburant 52 circule ainsi vers l'aval sur la surface extérieure de la base 50 et de l'extrémité évasée 48 du corps central 46. A la sortie de ce corps, grâce à la forme divergente, le film 52 est atomisé ce qui lui permet d'accrocher la flamme 54 située à l'intérieur de la chambre. De plus, la recirculation formée en bout de cette partie divergente 48 permet de stabiliser la flamme et d'augmenter ainsi les performances en extinction du foyer. En outre, l'extrémité évasée 42 comporte une rangée annulaire de trous 66 d'introduction d'air dans le foyer de combustion. Ces trous se situent à proximité d'une bride de fixation (non représentée) permettant de fixer le bol sur le fond de chambre 18, dans l'ouverture associée 32. In known manner, the injector 20 cooperates with the injection system 30 so that a film of fuel travels along the central body 46, downstream. The fuel film 52 thus flows downstream on the outer surface of the base 50 and the flared end 48 of the central body 46. At the exit of this body, thanks to the divergent shape, the film 52 is atomized. which allows it to hang the flame 54 located inside the chamber. In addition, the recirculation formed at the end of this diverging portion 48 stabilizes the flame and thus increase the extinction performance of the hearth. In addition, the flared end 42 has an annular row of holes 66 for introducing air into the combustion chamber. These holes are located near a fixing flange (not shown) for fixing the bowl on the chamber floor 18, in the associated opening 32.
La conception retenue met donc en œuvre une circulation d'un film de carburant 52 le long du corps central 46 du système d'injection, comme cela est par exemple de l'art antérieur. Cette conception à injection de film de carburant diffère de la conception dite à « spray » dans laquelle le carburant est injecté via une vrille, permettant également le passage de l'air. Du fait du passage du carburant dans la vrille, sous forme de spray, la perméabilité à l'air du système d'injection se trouve modifiée. En revanche, dans l'invention, l'air est destiné à circuler au travers du système d'injection 30 via un canal de traversée d'air 56 délimité entre le bol 40 et le corps central 46. Cette circulation, initiée de préférence par une vrille (non représentée sur les figures 3a et 3b), n'est pas perturbée par le film de carburant 48 circulant seulement sur la paroi intérieure de ce canal 56.  The selected design thus implements a circulation of a fuel film 52 along the central body 46 of the injection system, as is for example the prior art. This fuel film injection design differs from the so-called "spray" design in which the fuel is injected via a spin, also allowing the passage of air. Due to the passage of fuel in the spin, in the form of a spray, the air permeability of the injection system is modified. In contrast, in the invention, the air is intended to circulate through the injection system 30 via an air passage channel 56 delimited between the bowl 40 and the central body 46. This circulation, preferably initiated by a twist (not shown in Figures 3a and 3b), is not disturbed by the fuel film 48 flowing only on the inner wall of the channel 56.
En particulier, au niveau des extrémités évasées 42, 48, il est défini un canal tronconique de traversée d'air 60 constituant la partie aval du canal 56 précité. Le canal tronconique 60 est centré sur l'axe 22 et présente une section transversale référencée SI sur la figure 3a. L'une des particularités de l'invention réside dans le fait que le système d'injection intègre un degré de liberté de mouvement permettant de faire varier la section transversale du canal tronconique 60, en fonction des besoins rencontrés.  In particular, at the flared ends 42, 48, there is defined a frustoconical air passage channel 60 constituting the downstream portion of said channel 56. The frustoconical channel 60 is centered on the axis 22 and has a cross section referenced SI in Figure 3a. One of the peculiarities of the invention resides in the fact that the injection system integrates a degree of freedom of movement making it possible to vary the cross section of the frustoconical channel 60, according to the needs met.
Plus précisément, le système d'injection 30 comporte des moyens de mise en mouvement 62, permettant un déplacement relatif entre les première et seconde extrémités évasées 42, 48, selon l'axe 22. Ces moyens 62 sont de type conventionnel, par exemple intégrant un moteur linéaire, ou un électroaimant. Ils sont commandés par l'unité 13, et permettent de mettre en mouvement le corps central 46 à l'intérieur du bol 40, ce dernier étant fixe par rapport au fond de chambre 18 et à l'injecteur 20. Aussi, en fonction de la position relative axiale entre les premières et seconde extrémités évasées 42, 48, la section transversale du canal 60 varie. Sur la figure 3b, cette section référencée S2 est plus faible que la section SI de la figure 3a, car le corps central 46 a été déplacé vers l'amont par les moyens 62. More specifically, the injection system 30 comprises moving means 62, allowing relative movement between the first and second flared ends 42, 48, along the axis 22. These means 62 are of type conventional, for example incorporating a linear motor, or an electromagnet. They are controlled by the unit 13, and make it possible to set the central body 46 in motion inside the bowl 40, the latter being fixed with respect to the chamber bottom 18 and to the injector 20. Also, depending on the axial relative position between the first and second flared ends 42, 48, the cross section of the channel 60 varies. In FIG. 3b, this section referenced S2 is smaller than the section S1 of FIG. 3a, because the central body 46 has been moved upstream by the means 62.
Le fait de pouvoir faire varier la section de passage du canal 60 permet d'influencer la richesse du mélange air-carburant, qui impacte directement la stabilité de ce mélange. Cette faculté de variation de la section de passage du canal de traversée d'air permet d'influencer l'atomisation du carburant, qui se produit en sortie de la seconde extrémité évasée du corps central. En effet, l'atomisation peut être caractérisée par le rapport des quantités de mouvements d'air et de carburant, et donc directement dépendante de la section de passage du canal de traversée d'air. Cette atomisation peut également varier en fonction de la longueur du film de carburant 52 épousant extérieurement le corps central 46, cette longueur étant plus importante dans la position de la figure 3a que dans la position de la figure 3b sur laquelle le corps central 46 est en retrait, vers l'amont.  The fact of being able to vary the passage section of the channel 60 makes it possible to influence the richness of the air-fuel mixture, which directly impacts the stability of this mixture. This faculty of variation of the passage section of the air passage channel makes it possible to influence the atomization of the fuel, which occurs at the outlet of the second flared end of the central body. Indeed, the atomization can be characterized by the ratio of the amounts of air and fuel movements, and therefore directly dependent on the passage section of the air passage channel. This atomization may also vary according to the length of the fuel film 52 conforming externally to the central body 46, this length being greater in the position of FIG. 3a than in the position of FIG. 3b on which the central body 46 is in position. withdrawal, upstream.
Avantageusement, le phénomène d'atomisation présente un impact direct sur la stabilité de la chambre de combustion, sur la capacité d'allumage au sol et de rallumage en altitude, ou encore sur les émissions polluantes en régime de ralenti. L'ensemble de ces paramètres peuvent ainsi être optimisés pour tous les points de fonctionnement de la turbomachine. Par exemple, une perte de charge importante de l'air à travers le système d'injection est avantageuse pour atomiser et mélanger le carburant au ralenti, afin d'augmenter la stabilité de la flamme. La position de la figure 3b, avec la section réduite S2, sera donc préférée pour ce régime de ralenti de la turbomachine. En revanche, pour limiter les pertes en termes de consommation spécifique, la position de la figure 3a sera préférentiellement retenue pour les régimes de plein gaz et de croisière. En référence à présent aux figures 4a et 4b, il est représenté le système d'injection 30 selon un mode de réalisation préféré de l'invention. Sur ces figures, les éléments portant les mêmes références numériques que des éléments des figures de principe 3a et 3b, correspondent à des éléments identiques ou similaires. Pour des raisons de clarté, les moyens de mise en mouvement 62 du corps principal 46 n'ont pas été représentés sur ces figures 4a et 4b. Néanmoins, ces moyens 62 sont bien évidemment prévus et commandés pour déplacer le corps principal 46 de la position de la figure 4a à celle de la figure 4b, et inversement. Advantageously, the atomization phenomenon has a direct impact on the stability of the combustion chamber, on the ground ignition and reignition capability at altitude, or on pollutant emissions at idle speed. All of these parameters can thus be optimized for all operating points of the turbomachine. For example, a significant pressure drop of air through the injection system is advantageous for atomizing and mixing the fuel at idle, in order to increase the stability of the flame. The position of FIG. 3b, with the reduced section S2, will therefore be preferred for this idling speed of the turbomachine. On the other hand, in order to limit the losses in terms of specific consumption, the position of FIG. 3a will preferably be retained for the full throttle and cruising regimes. Referring now to Figures 4a and 4b, there is shown the injection system 30 according to a preferred embodiment of the invention. In these figures, the elements bearing the same reference numbers as elements of the principle figures 3a and 3b correspond to identical or similar elements. For the sake of clarity, the moving means 62 of the main body 46 have not been shown in these figures 4a and 4b. Nevertheless, these means 62 are obviously provided and controlled to move the main body 46 from the position of Figure 4a to that of Figure 4b, and vice versa.
Aussi, dans ce mode de réalisation préféré, le système d'injection 30 comporte une structure intermédiaire 70, agencée radialement entre la base 50 du corps central 46 et la base 44 du bol 40. C'est relativement à cette structure intermédiaire 70 que le corps principal 46 est capable d'être déplacé axialement entre les deux positions des figures 4a et 4b, la structure 70 restant quant à elle fixe relativement à l'injecteur 20 et le bol 40.  Also, in this preferred embodiment, the injection system 30 comprises an intermediate structure 70, arranged radially between the base 50 of the central body 46 and the base 44 of the bowl 40. It is relatively to this intermediate structure 70 that the main body 46 is capable of being displaced axially between the two positions of FIGS. 4a and 4b, while the structure 70 remains fixed relative to the injector 20 and the bowl 40.
La structure intermédiaire 70 délimite avec la surface extérieure de la base 50 un canal annulaire axial 72 d'écoulement du film de carburant 52, en direction de la seconde extrémité évasée 48 du corps central 46. C'est en effet ce canal 72 qui est alimenté de façon connue par l'injecteur 20 et qui permet de générer le film de carburant 52 de faible épaisseur le long du corps central 46, avant de rencontrer l'air introduit dans le système d'injection. Dans le mode de réalisation préféré représenté, le carburant contourne l'extrémité amont du corps central plein 46 avant d'épouser la paroi extérieure de celui-ci délimitant intérieurement le canal annulaire axial 72.  The intermediate structure 70 defines with the outer surface of the base 50 an axial annular channel 72 for the flow of the fuel film 52, in the direction of the second flared end 48 of the central body 46. It is indeed this channel 72 which is fed in known manner by the injector 20 and which allows to generate the thin film of fuel 52 along the central body 46, before encountering the air introduced into the injection system. In the preferred embodiment shown, the fuel bypasses the upstream end of the solid central body 46 before marrying the outer wall thereof internally defining the axial annular channel 72.
La base 44 du bol 40 comporte ici deux parois concentriques 44a, 44b entre lesquelles est agencée une vrille 76 d'introduction d'air entre les deux parois concentriques, cette vrille étant à caractère axial ou radial. L'air provenant de la vrille 76 et circulant entre les deux parois externe et interne 44a, 44b, rejoint ensuite le canal tronconique 60 dans lequel le film de carburant 52 lèche la surface extérieure de l'extrémité tronconique 48 du corps central 46. La paroi interne 44b entoure la structure intermédiaire 70, de manière à délimiter entre elles un canal 80 débouchant vers l'aval. Dans ce mode de réalisation préféré, le canal 80 n'est pas destiné à être traversé par un flux d'air. The base 44 of the bowl 40 here comprises two concentric walls 44a, 44b between which is arranged a swirl 76 for introducing air between the two concentric walls, this swirl being of axial or radial character. The air coming from the auger 76 and circulating between the two outer and inner walls 44a, 44b, then joins the frustoconical channel 60 in which the fuel film 52 licks the outer surface of the frustoconical end 48 of the central body 46. internal wall 44b surrounds the intermediate structure 70, so as to delimit between them a channel 80 opening downstream. In this preferred embodiment, the channel 80 is not intended to be traversed by an air flow.
Dans le canal 60 de plus grande largeur que celle du canal 72 dans lequel le film de carburant 52 est créé, ce dernier reste confiné le long de la surface latérale de l'extrémité tronconique 48 du corps central 46, grâce au passage de l'air 82 dans ce même canal 60.  In the channel 60 of greater width than that of the channel 72 in which the fuel film 52 is created, the latter remains confined along the lateral surface of the frustoconical end 48 of the central body 46, thanks to the passage of the air 82 in the same channel 60.
A cet égard, il est noté que la partie grisée 52 la plus foncée de la figure 5 représente le trajet du carburant depuis l'injecteur 20 jusqu'à la flamme 54, tandis que la partie grisée la plus claire 82 représente le flux d'air.  In this regard, it is noted that the darkest shaded portion 52 of Figure 5 represents the fuel path from the injector 20 to the flame 54, while the lightest gray portion 82 represents the flow of the fuel. air.
Bien entendu, diverses modifications peuvent être apportées par l'homme du métier à l'invention qui vient d'être décrite sans sortir du cadre de l'exposé de l'invention.  Of course, various modifications may be made by those skilled in the art to the invention which has just been described without departing from the scope of the disclosure of the invention.

Claims

REVENDICATIONS
1. Ensemble comprenant un système d'injection (30) pour chambre de combustion (8) de turbomachine d'aéronef ainsi qu'un injecteur de carburant (20) coopérant avec ledit système d'injection (30), ce dernier comprenant un bol aérodynamique (40) fixe par rapport à l'injecteur (20) et comportant une première extrémité évasée vers l'aval (42) et centrée sur un axe central (22) du système d'injection, celui-ci comportant également un corps central (46) le long duquel un film de carburant (52) est destiné à circuler vers l'aval, 1. An assembly comprising an injection system (30) for an aircraft turbomachine combustion chamber (8) and a fuel injector (20) cooperating with said injection system (30), the latter comprising a bowl aerodynamic device (40) fixed relative to the injector (20) and having a first end flared downstream (42) and centered on a central axis (22) of the injection system, the latter also comprising a central body (46) along which a fuel film (52) is intended to flow downstream,
caractérisé en ce que le corps central (46) comporte une seconde extrémité évasée vers l'aval (48) et centrée sur l'axe central (22) du système d'injection, lesdites première et seconde extrémités évasées (42, 48) délimitant entre elles un canal de traversée d'air (60), et en ce que le système (30) comporte des moyens de mise en mouvement (62) permettant un déplacement relatif entre lesdites première et seconde extrémités évasées (42, 48), selon l'axe central (22) du système d'injection, par mise en mouvement du corps central (46) relativement à l'injecteur (20).  characterized in that the central body (46) has a second end flared downstream (48) and centered on the central axis (22) of the injection system, said first and second flared ends (42, 48) defining between them an air passage channel (60), and in that the system (30) comprises moving means (62) for relative movement between said first and second flared ends (42, 48), the central axis (22) of the injection system, by moving the central body (46) relative to the injector (20).
2. Ensemble selon la revendication 1, caractérisé en ce que lesdites première et seconde extrémités évasées (42, 48) sont de forme tronconique et en ce que qu'elles délimitent entre elles un canal tronconique de traversée d'air (60), de section variable en fonction d'une position relative axiale entre lesdites première et seconde extrémités évasées (42, 48). 2. An assembly according to claim 1, characterized in that said first and second flared ends (42, 48) are of frustoconical shape and in that they delimit between them a truncated cone channel air (60), variable section according to an axial relative position between said first and second flared ends (42, 48).
3. Ensemble selon la revendication 1 ou la revendication 2, caractérisé en ce que le système d'injection comporte une structure intermédiaire (70) agencée radialement entre une base (50) du corps central (46) et une base (44) du bol aérodynamique (40), ladite structure intermédiaire (70) délimitant avec la base (50) du corps central (46) un canal axial (72) d'écoulement du film de carburant (52) en direction de ladite seconde extrémité évasée (48) du corps central. 3. An assembly according to claim 1 or claim 2, characterized in that the injection system comprises an intermediate structure (70) arranged radially between a base (50) of the central body (46) and a base (44) of the bowl. aerodynamic device (40), said intermediate structure (70) defining with the base (50) of the central body (46) an axial flow channel (72) of the fuel film (52) towards said second flared end (48) of the central body.
4. Ensemble selon la revendication 3, caractérisé en ce que la base (44) du bol aérodynamique (40) comporte deux parois concentriques (44a, 44b) entre lesquelles est agencée une vrille (76) d'introduction d'air entre les deux parois concentriques. 4. An assembly according to claim 3, characterized in that the base (44) of the aerodynamic bowl (40) comprises two concentric walls (44a, 44b) between which is arranged a swirl (76) of introduction of air between the two concentric walls.
5. Ensemble selon l'une quelconque des revendications précédentes, caractérisé en ce que les moyens de mise en mouvement (62) comprennent un moteur. 5. An assembly according to any one of the preceding claims, characterized in that the moving means (62) comprise a motor.
6. Ensemble selon l'une quelconque des revendications précédentes, caractérisé en ce que ladite première extrémité évasée (42) du bol aérodynamique (40) est percée de trous (66) d'introduction d'air dans un foyer de combustion délimité par ce bol. 6. An assembly according to any one of the preceding claims, characterized in that said first flared end (42) of the aerodynamic bolt (40) is pierced with holes (66) of introduction of air into a combustion chamber defined by this bowl.
7. Chambre de combustion (8) de turbomachine d'aéronef comprenant un fond de chambre (18) percé d'ouvertures (32) espacées les unes des autres, la chambre de combustion comprenant, associé à chaque ouverture (32) du fond de chambre (18), un ensemble selon l'une quelconque des revendications précédentes. 7. Aircraft turbomachine combustion chamber (8) comprising a chamber bottom (18) pierced with apertures (32) spaced from each other, the combustion chamber comprising, associated with each opening (32) of the bottom of chamber (18), an assembly according to any one of the preceding claims.
8. Turbomachine (1) d'aéronef comprenant une chambre de combustion (8) selon la revendication précédente. 8. Turbine engine (1) comprising a combustion chamber (8) according to the preceding claim.
EP16703587.2A 2015-01-20 2016-01-20 Ensemble comprising a fuel injection system for an aircraft turbomachine combustion chamber and a fuel injector Active EP3247945B1 (en)

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FR1550444A FR3031798B1 (en) 2015-01-20 2015-01-20 FUEL INJECTION SYSTEM FOR AIRCRAFT TURBINE ENGINE COMPRISING A VARIABLE SECTION AIR AIR CHANNEL
PCT/FR2016/050107 WO2016116700A1 (en) 2015-01-20 2016-01-20 Fuel injection system for aircraft turbomachine, comprising a variable section air through duct

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Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2930191A (en) * 1953-01-29 1960-03-29 Phillips Petroleum Co Air-fuel control in prevaporizer type combustion chambers
US3490230A (en) * 1968-03-22 1970-01-20 Us Navy Combustion air control shutter
US4150539A (en) * 1976-02-05 1979-04-24 Avco Corporation Low pollution combustor
US4271675A (en) * 1977-10-21 1981-06-09 Rolls-Royce Limited Combustion apparatus for gas turbine engines
DE3518080A1 (en) * 1985-05-20 1986-11-20 Stubinen Utveckling AB, Stockholm METHOD AND DEVICE FOR BURNING LIQUID AND / OR SOLID FUELS IN POWDERED FORM
US5235813A (en) * 1990-12-24 1993-08-17 United Technologies Corporation Mechanism for controlling the rate of mixing in combusting flows
FR2685452B1 (en) 1991-12-24 1994-02-11 Snecma FUEL INJECTION DEVICE FOR A TURBOMACHINE COMBUSTION CHAMBER.
US5217363A (en) * 1992-06-03 1993-06-08 Gaz Metropolitan & Co., Ltd. And Partnership Air-cooled oxygen gas burner assembly
US6199367B1 (en) * 1996-04-26 2001-03-13 General Electric Company Air modulated carburetor with axially moveable fuel injector tip and swirler assembly responsive to fuel pressure
FR2832493B1 (en) 2001-11-21 2004-07-09 Snecma Moteurs MULTI-STAGE INJECTION SYSTEM OF AN AIR / FUEL MIXTURE IN A TURBOMACHINE COMBUSTION CHAMBER
FR2875585B1 (en) 2004-09-23 2006-12-08 Snecma Moteurs Sa AERODYNAMIC SYSTEM WITH AIR / FUEL INJECTION EFFERVESCENCE IN A TURBOMACHINE COMBUSTION CHAMBER
US20100031669A1 (en) * 2008-08-06 2010-02-11 Cessna Aircraft Company Free Turbine Generator For Aircraft

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FR3031798A1 (en) 2016-07-22
WO2016116700A1 (en) 2016-07-28
FR3031798B1 (en) 2018-08-10
US10371384B2 (en) 2019-08-06
US20180010799A1 (en) 2018-01-11
EP3247945B1 (en) 2018-12-05

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