EP3247945B1 - Ensemble mit einem kraftstoffeinspritzsystem für eine flugzeugturbomaschinenbrennkammer und einem brennstoffinjektor - Google Patents

Ensemble mit einem kraftstoffeinspritzsystem für eine flugzeugturbomaschinenbrennkammer und einem brennstoffinjektor Download PDF

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Publication number
EP3247945B1
EP3247945B1 EP16703587.2A EP16703587A EP3247945B1 EP 3247945 B1 EP3247945 B1 EP 3247945B1 EP 16703587 A EP16703587 A EP 16703587A EP 3247945 B1 EP3247945 B1 EP 3247945B1
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EP
European Patent Office
Prior art keywords
injection system
central body
fuel
combustion chamber
widening
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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.)
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Application number
EP16703587.2A
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English (en)
French (fr)
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EP3247945A1 (de
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
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Safran Aircraft Engines SAS
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Publication of EP3247945A1 publication Critical patent/EP3247945A1/de
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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 sought to limit idle pollution as much as possible, which is closely related to the capacity of the injection system to atomize and mix the fuel injected with the air.
  • Such aerodynamic injection systems are known from the documents FR 2,875,585 , or documents FR 2,685,452 and FR 2,832,493 .
  • obtaining optimal performance for certain points of operation of the turbomachine goes through design 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.
  • 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.
  • an aircraft turbine engine 1 there is shown an aircraft turbine engine 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.
  • 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 on the figure 2 . It has in particular an outer shell 14 centered on the axis 3, an inner shell 16 also centered on the same axis, and a chamber bottom 18 connecting the two rings at their upstream end.
  • Fuel injectors 20 are regularly distributed over the chamber bottom, in the circumferential direction (a single injector being visible on the figure 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, represented schematically on the figure 2 .
  • the injection system 30 cooperates upstream with the nozzle nose 21, while it opens downstream into the combustion chamber 8.
  • the injection system 30 is housed in an opening 32 made 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 therefore uses 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 spin (not shown on the Figures 3a and 3b ), is not disturbed by the fuel film 48 flowing only on the inner wall of this 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 S1 on the 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. On the figure 3b , this section referenced S2 is smaller than the section S1 of the figure 3a because the central body 46 has been moved upstream by the means 62.
  • 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 the figure 3a that in the position of the figure 3b on which the central body 46 is recessed, 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 the figure 3b with the reduced section S2, will be preferred for this idle speed of the turbomachine.
  • the position of the figure 3a will be preferentially retained for the full gas and cruise plans.
  • FIG. 4a and 4b it is represented 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 represented on these Figures 4a and 4b . Nevertheless, these means 62 are obviously provided and controlled to move the main body 46 of the position of the figure 4a to that of the 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 able to be moved axially between the two positions of the Figures 4a and 4b , the structure 70 remaining fixed in relation 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 gray portion 52 of the figure 5 represents the fuel path from the injector 20 to the flame 54, while the lightest gray portion 82 represents the air flow.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Spray-Type Burners (AREA)
  • Fuel-Injection Apparatus (AREA)

Claims (8)

  1. Gesamtanordnung mit einem Einspritzsystem (30) für Brennkammern (8) von Luftfahrzeug-Turbotriebwerken sowie einem Brennstoff-Injektor (20), der mit diesem Einspritzsystem (30) zusammenwirkt, wobei letzteres einen stromlinienförmigen Mantel (40) aufweist, der feststehend gegenüber dem Injektor (20) ist und ein erstes aufgeweitetes Ende (42) aufweist, das in Strömungsrichtung abwärts gerichtet ist und auf einer zentralen Achse (22) des Einspritzsystems (30) zentriert ist,
    wobei dieses ferner einen zentralen Körper (46) aufweist, an dem entlang ein Brennstofffilm (52) in Strömungsrichtung abwärts laufen soll,
    wobei der zentrale Körper (46) ein zweites aufgeweitetes Ende (48) aufweist, das in Strömungsrichtung abwärts gerichtet ist und auf der zentralen Achse (22) des Einspritzsystems (30) zentriert ist,
    wobei diese beiden aufgeweiteten Enden (42, 48) zwischen sich einen Luftströmungskanal (60) bilden,
    wobei das Einspritzsystem (30) Bewegungsmittel (62) aufweist, die ein relatives Verschieben der genannten beiden aufgeweiteten Enden (42, 48) zueinander auf der zentralen Achse (22) des Einspritzsystems (30) durch In-Bewegung-Setzen des zentralen Körpers (46) gegenüber dem Injektor (20) ermöglichen.
  2. Gesamtanordnung nach Anspruch 1,
    dadurch gekennzeichnet,
    dass die genannten beiden aufgeweiteten Enden (42, 48) kegelstumpfförmig sind und dass sie zwischen sich einen kegelstumpfförmigen Luftströmungskanal (60) bilden, dessen Querschnitt in Abhängigkeit von einer relativen axialen Position der beiden aufgeweiteten Enden (42, 48) zueinander variabel ist.
  3. Gesamtanordnung nach Anspruch 1 oder Anspruch 2,
    dadurch gekennzeichnet,
    dass das Einspritzsystem (30) ein zwischenliegendes Bauteil (70) aufweist, das radial zwischen einer Basis (50) des zentralen Körpers (46) und einer Basis (44) des stromlinienförmigen Mantels (40) angebracht ist, wobei dieses zwischenliegende Bauteil (70) zusammen mit der Basis (50) des zentralen Körpers (46) einen axialen Kanal (72) zum Fließen des Brennstofffilms (52) in Richtung des zweiten aufgeweiteten Endes (48) des zentralen Körpers (46) umgrenzt.
  4. Gesamtanordnung nach Anspruch 3,
    dadurch gekennzeichnet,
    dass die Basis (44) des stromlinienförmigen Mantels (40) zwei konzentrische Wandungen (44a, 44b) aufweist, zwischen denen eine Schnecke (76) zur Zuführung von Luft zwischen die beiden konzentrischen Wandungen (44a, 44b) angeordnet ist.
  5. Gesamtanordnung nach einem der vorherigen Ansprüche,
    dadurch gekennzeichnet,
    dass die Bewegungsmittel (62) einen Motor umfassen.
  6. Gesamtanordnung nach einem der vorherigen Ansprüche,
    dadurch gekennzeichnet,
    dass in das erste aufgeweitete Ende (42) des stromlinienförmigen Mantels (40) Löcher (66) zur Zuführung von Luft in einen von diesem stromlinienförmigen Mantel (40) umgrenzten Brennraum gebohrt sind.
  7. Brennkammer (8) von Luftfahrzeug-Turbotriebwerken mit einem Kammerboden (18), durch den Öffnungen (32) in Abstand voneinander gebohrt sind, wobei die Brennkammer (8) in Verbindung mit jeder Öffnung (32) des Kammerbodens (18) eine Gesamtanordnung nach einem der vorherigen Ansprüche aufweist.
  8. Turbotriebwerk (1) für Luftfahrzeuge, das eine Brennkammer (8) nach dem vorherigen Anspruch aufweist.
EP16703587.2A 2015-01-20 2016-01-20 Ensemble mit einem kraftstoffeinspritzsystem für eine flugzeugturbomaschinenbrennkammer und einem brennstoffinjektor Active EP3247945B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1550444A FR3031798B1 (fr) 2015-01-20 2015-01-20 Systeme d'injection de carburant pour turbomachine d'aeronef, comprenant un canal de traversee d'air a section variable
PCT/FR2016/050107 WO2016116700A1 (fr) 2015-01-20 2016-01-20 Systeme d'injection de carburant pour turbomachine d'aeronef, comprenant un canal de traversee d'air a section variable

Publications (2)

Publication Number Publication Date
EP3247945A1 EP3247945A1 (de) 2017-11-29
EP3247945B1 true EP3247945B1 (de) 2018-12-05

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EP16703587.2A Active EP3247945B1 (de) 2015-01-20 2016-01-20 Ensemble mit einem kraftstoffeinspritzsystem für eine flugzeugturbomaschinenbrennkammer und einem brennstoffinjektor

Country Status (4)

Country Link
US (1) US10371384B2 (de)
EP (1) EP3247945B1 (de)
FR (1) FR3031798B1 (de)
WO (1) WO2016116700A1 (de)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3096114B1 (fr) 2019-05-13 2022-10-28 Safran Aircraft Engines Chambre de combustion comprenant des moyens de refroidissement d’une zone d’enveloppe annulaire en aval d’une cheminée

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 (de) * 1985-05-20 1986-11-20 Stubinen Utveckling AB, Stockholm Verfahren und vorrichtung zum verbrennen fluessiger und/oder fester brennstoffe in pulverisierter form
US5235813A (en) * 1990-12-24 1993-08-17 United Technologies Corporation Mechanism for controlling the rate of mixing in combusting flows
FR2685452B1 (fr) 1991-12-24 1994-02-11 Snecma Dispositif d'injection de carburant pour une chambre de combustion de turbomachine.
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 (fr) 2001-11-21 2004-07-09 Snecma Moteurs Systeme d'injection multi-etages d'un melange air/carburant dans une chambre de combustion de turbomachine
FR2875585B1 (fr) 2004-09-23 2006-12-08 Snecma Moteurs Sa Systeme aerodynamique a effervescence d'injection air/carburant dans une chambre de combustion de turbomachine
US20100031669A1 (en) * 2008-08-06 2010-02-11 Cessna Aircraft Company Free Turbine Generator For Aircraft

Non-Patent Citations (1)

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Title
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Publication number Publication date
EP3247945A1 (de) 2017-11-29
WO2016116700A1 (fr) 2016-07-28
FR3031798B1 (fr) 2018-08-10
US20180010799A1 (en) 2018-01-11
US10371384B2 (en) 2019-08-06
FR3031798A1 (fr) 2016-07-22

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