EP3568638B1 - Brennkammer für einen turbinenmotor - Google Patents

Brennkammer für einen turbinenmotor Download PDF

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Publication number
EP3568638B1
EP3568638B1 EP18700941.0A EP18700941A EP3568638B1 EP 3568638 B1 EP3568638 B1 EP 3568638B1 EP 18700941 A EP18700941 A EP 18700941A EP 3568638 B1 EP3568638 B1 EP 3568638B1
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EP
European Patent Office
Prior art keywords
radially
sealing member
combustion chamber
cylindrical part
annular
Prior art date
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EP18700941.0A
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English (en)
French (fr)
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EP3568638A1 (de
Inventor
Jacques Marcel Arthur Bunel
Dan Ranjiv JOORY
Patrice André Commaret
Romain Nicolas Lunel
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Safran Aircraft Engines SAS
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Safran Aircraft Engines SAS
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Publication of EP3568638A1 publication Critical patent/EP3568638A1/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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/60—Support structures; Attaching or mounting 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
    • F23R3/002—Wall structures
    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00012—Details of sealing devices

Definitions

  • the present invention relates to a combustion chamber for a turbomachine, in particular for an aircraft turbojet or turboprop.
  • a turbomachine in particular a double-body turbomachine, conventionally comprises, from upstream to downstream, a fan, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine and a low-pressure turbine.
  • upstream and downstream are defined with respect to the direction of air circulation in the turbomachine.
  • internal and exitterior are defined radially with respect to the axis of the turbomachine.
  • a combustion chamber conventionally comprises a radially outer annular shell, a radially inner annular shell, coaxial with the radially outer shell, and a bottom wall connecting the radially outer shell and the radially inner shell.
  • the bottom wall has radially outer and inner cylindrical parts. Furthermore, the outer and inner shrouds each have a cylindrical part at their upstream end.
  • the outer cylindrical part of the bottom wall is fixed by bolting to the cylindrical part of the outer shell.
  • the internal cylindrical part of the bottom wall is fixed by bolting to the cylindrical part of the internal ferrule.
  • a radial annular clearance exists between the aforementioned cylindrical parts.
  • the interface between the cylindrical parts of the ferrules and the cylindrical parts of the bottom wall delimits lobes.
  • Such lobes between the aforementioned cylindrical parts allow the entry of parasitic air into the combustion chamber or the exit of combustion gas out of said chamber. This penalizes the efficiency of the combustion chamber and can generate a pollution phenomenon.
  • these openings may represent an air passage surface area of approximately 300 mm 2 , ie 3% of the total air flow rate entering the combustion chamber.
  • the object of the invention is in particular to provide a simple, effective and economical solution to this problem.
  • the combustion chamber may comprise a second annular sealing member, coaxial with said radially inner and outer shrouds, the second sealing member being interposed radially between the bottom wall and the radially inner shroud.
  • the sealing member makes it possible to fill in the radial play between the bottom wall and the corresponding ferrule of the combustion chamber, in order to limit the passage of air to the aforementioned interface zones. This improves the performance of the turbomachine and limits the sources of pollution.
  • Each sealing member can be sectored and include at least two angular sectors.
  • each angular sector can be deformed slightly so as to adapt to the real diameter of the interface zone considered.
  • Each angular sector can then optimally close off said interface zone.
  • the angular sectors can be distributed over the circumference with a total angular play between them of between 0 and 1 ° or 0 and 5 mm.
  • the total clearance between the sectors is for example between 0 and 1 ° or 0 and 5 mm, for an implantation diameter of between 500 and 650 mm, for example.
  • This spacing makes it possible in particular to allow the deformations of the sectors during their shaping to the aforementioned interface zones.
  • the outer shell of the combustion chamber may include a cylindrical portion surrounding a radially outer cylindrical portion of the bottom wall, the bottom wall may further include at least one radially inner cylindrical portion surrounding a cylindrical portion of the inner shell of the chamber.
  • combustion chamber the first sealing member being able to be interposed between the cylindrical part of the outer shell and the outer cylindrical part of the bottom wall
  • the second sealing member being able to be interposed between the cylindrical part of the inner shell and the internal cylindrical part of the bottom wall.
  • the thermal protection member makes it possible to protect the bottom wall and the elements situated upstream of the latter from high temperatures within the combustion chamber.
  • the radially outer rim of the thermal protection member may be located near the outer shell of the chamber, that is to say at a distance of between 0.1 and 2.5 mm.
  • the radially internal rim of the thermal protection member may be located near the internal ferrule of the chamber, that is to say at a distance of between 0.1 and 2.5 mm.
  • Each sealing member can be made of a nickel-based alloy, for example of the Hastelloy® type, or of a cobalt-based alloy.
  • Such a material is able to withstand thermal stresses in operation.
  • Each sealing member may have a thickness of between 0.8 and 3 mm.
  • the sealing member can be provided with fixing holes, evenly distributed around the circumference.
  • Each sealing member can be fixed to the bottom wall by means of fixing means, such as screws.
  • Said screws or said rivets can be engaged in the fixing holes of the corresponding sealing member.
  • the screws or rivets can first be engaged in the holes located at the level of the circumferentially median zone of the corresponding angular sector, then gradually in the holes located near the circumferential ends of the sector.
  • Each sector of the sealing member may be in the form of a strip in the form of an arc of a circle.
  • the thermal protection member may be in the form of a sheet whose thickness is between 0.5 and 1.5 mm.
  • the thermal protection member may be made of a nickel-based alloy, for example of the Hastelloy® type, or of a cobalt-based alloy.
  • At least one of the internal and external ferrules of the combustion chamber may have cutouts opening upstream.
  • the combustion chamber may include an upstream cover comprising a radially outer annular fixing part, fixed to the outer shell of the combustion chamber, said cover further comprising a radially inner annular fixing part, fixed to the outer shell of the chamber. combustion.
  • the radially internal surface of the outer shell may include an annular recess, the downstream axial end of which forms an annular radial shoulder, the first sealing member being housed, at least in part, in the recess, the downstream end of each sector angular of the first sealing member being able to come to rest on the shoulder.
  • the invention also provides a turbomachine, in particular an aircraft turbojet or turboprop, comprising a combustion chamber of the aforementioned type.
  • turbomachine 1 is of the double-body and double-flow type, and extends along a longitudinal axis X.
  • the turbomachine 1 comprises a fan 2 drawing in an air flow which is divided into a primary flow and a secondary flow.
  • the primary flow passes through a primary stream 3 which successively comprises, from upstream AM to downstream AV, a low pressure compressor 4 and a high pressure compressor 5.
  • the air is injected and mixed with a fuel. in a combustion chamber 6.
  • hot gases pass successively through a high pressure turbine 7 and a low pressure turbine 8 before being ejected from the turbomachine 1 by an ejection nozzle 9.
  • the secondary flow passes through a secondary vein 10 surrounding the primary vein 3.
  • the figures 2 to 7 show several embodiments of the combustion chamber 6 of the turbomachine 1 according to the invention.
  • the combustion chamber 6 comprises a radially outer annular shell 11, a radially inner annular shell 12, and an annular bottom wall 13 extending radially and connecting the radially outer shell 11 and the radially inner shell 12.
  • the outer shell 11 has a general frustoconical shape widening downstream AV.
  • the outer shell 11 comprises, at its upstream end, a cylindrical part 14.
  • Said cylindrical part 14 has holes distributed over the circumference.
  • the cylindrical part 14 further comprises cutouts 15 distributed over the circumference, said cutouts 15 opening upstream AM.
  • the outer shell 11 further comprises air inlet holes 16, also called primary holes.
  • the internal ferrule 12 has a general frustoconical shape widening towards the downstream AV.
  • the internal ferrule 12 comprises, at its upstream end, a cylindrical part 17.
  • Said cylindrical part 17 has holes distributed over the circumference.
  • the cylindrical part 17 further comprises cutouts distributed over the circumference, said cutouts opening upstream AM.
  • the internal ferrule 12 further comprises air inlet holes 18.
  • the bottom wall 13 is annular and comprises a part 19 of generally frustoconical shape or extending radially.
  • the radially outer periphery of the frustoconical or radial part is extended by a cylindrical part 20 extending upstream AM.
  • the radially inner periphery of the frustoconical or radial part is extended by a cylindrical part 21 extending upstream AM.
  • the bottom wall 13 has openings 22 distributed over the circumference of the part Frustoconical 19.
  • the cylindrical parts 20, 21 of the bottom wall 13 have fixing holes 23 distributed over the circumference.
  • the combustion chamber 6 further comprises an annular cover 24 of section in the general shape of a C in cross section, located upstream AM of the bottom wall 13.
  • the radially outer periphery of the cover 24 comprises a cylindrical part 25.
  • the radially inner periphery of the cover 24 comprises a cylindrical part 26.
  • the radially median zone 27 of the cover 24 has openings 28 located axially facing the openings 22 of the bottom wall 13.
  • the outer cylindrical part 25 of the cover 24, the cylindrical part 14 of the outer shell 11 and the outer cylindrical part 20 of the bottom wall 13 are fixed to each other by means of bolts 29 distributed around the circumference and engaged in the holes of the cylindrical part 14 of the outer shell and the fixing holes 23 of the bottom wall 13. More particularly, the outer cylindrical part 25 of the cover 24 surrounds the cylindrical part 14 of the outer shell 11, which itself surrounds the outer cylindrical part 20 of the bottom wall 13.
  • the internal cylindrical part 26 of the cover 24, the cylindrical part 17 of the internal ferrule 12 and the internal cylindrical part 21 of the bottom wall 13 are fixed to each other by means of bolts 30 distributed around the circumference and engaged in the holes in the cylindrical part 17 of the inner shell 12 and the fixing holes 23 in the bottom wall 13. More particularly, the inner cylindrical part 21 of the bottom wall 13 surrounds the cylindrical part 17 of the inner shell 12, which itself surrounds the internal cylindrical part 26 of the cover 24.
  • Each opening 22 of the bottom wall 13 is used for mounting a fuel injection device 31.
  • the fuel injection device 31 is connected to an injection pipe 32 forming a fuel supply pipe, said fuel injection device.
  • injection nozzle 32 passing through the corresponding opening 28 of the cover 24.
  • the structure of the injection device 31 is known per se and will not be described in more detail.
  • the downstream end (not shown) of the combustion chamber 6 is fixed to an outer casing 33.
  • Said outer casing 33 comprises a radially outer wall 34 and a radially inner wall 35 connected at their upstream end.
  • the junction 36 between the radially inner wall and the outer wall comprises an air inlet port 37, allowing the air from the high pressure compressor 5 to enter the internal volume of the outer casing 33.
  • the air passes through thus said orifice 37 then divides into a first part which passes through the opening 28 of the cover 24 and enters the fuel injection device 31 in which it is mixed with the fuel.
  • a second part of the air bypasses the cover 24 then enters the combustion chamber 6 through the holes 16, 18 of the inner 12 and outer 11 ferrules.
  • the casing is formed in one piece, that is to say that the radially outer 34 and radially inner 35 walls form a single piece with the junction 36.
  • the walls 34 35 and the junction 36 are, for example, integral.
  • the walls 34, 35 could be attached and fixed to the junction 36, the walls 34, 35 and the junction 36 then being independent of each other.
  • a radial annular clearance exists between the aforementioned cylindrical parts 14, 17, 20, 21 of the shrouds 11, 12 and of the bottom wall 13, in order to allow the bottom wall 13 to be fitted between the shrouds 11, 12 and due to dimensional manufacturing tolerances.
  • the combustion chamber 6 comprises first and second annular sealing members 38a, 38b aimed at filling these clearances.
  • the first sealing member 38a is interposed radially between the bottom wall 13 and the radially outer shell 11.
  • the second sealing member 38b is interposed radially between the bottom wall 13 and the radially inner shell 12.
  • first sealing member 38a and the second sealing member 38b are concentric and of identical structures.
  • Each annular sealing member 38a, 38b is annular and formed of at least two angular sectors 39a, 39b (only the first sealing member 38a is shown in figure 4 ), here two angular sectors 39a, 39b.
  • Each sector 39a, 39b is curved and has the shape of a circular arc.
  • Each sector 38a, 38b comprises, on its circumference, fixing holes 40 regularly distributed around the circumference.
  • the angular sectors 39a, 39b are distributed over the circumference of the cylindrical part 20 of the bottom wall 13 and are spaced slightly by a clearance noted j with respect to each other, at their ends, like this is better visible at the figure 5 .
  • the total angular play between the sectors is for example between 0 and 1 ° or 0 and 5 mm.
  • Each sector 39a, 39b of each sealing member 38a, 38b is for example made of a nickel-based alloy, for example of the Hastelloy® type, or of a cobalt-based alloy.
  • Each sector 38a, 38b has for example a thickness of between 0.8 and 3 mm.
  • each sealing member 38a, 38b are fixed by bolts (not shown) engaged only in some of the fixing holes 23 of the corresponding cylindrical part 20, 21 of the bottom wall 13 and in the holes of the sectors 39a, 39b of the sealing member 38a, 38b.
  • the screw heads or the nuts of these bolts are located at the level of the cutouts 15 of the corresponding ferrule 11, 12.
  • the combustion chamber 6 further comprises a thermal protection member 41 located downstream AV of the bottom wall 13, in the form of an annular sheet.
  • the protection member 41 comprises an annular portion 42, of frustoconical shape or extending in a radial plane, the inner and outer peripheral edges of which are extended by annular flanges 43, 44 extending axially upstream AM ( figure 3 ).
  • the outer rim 43 of the protective member 41 is interposed radially between the cylindrical part 14 of the outer shell 11 and the outer cylindrical part 20 of the bottom wall 13. Furthermore, the outer rim 43 of the protective member 41 is located downstream AV of the first sealing member 38a.
  • the internal rim of the protective member 41 (not shown on the figure 3 ) is interposed radially between the cylindrical part 17 of the inner shell 12 and the inner cylindrical part 21 of the bottom wall 13. Furthermore, the inner rim of the protective member 41 is located downstream AV of the second member of ' sealing 38b.
  • the flanges 43, 44 of the protective member may extend axially downstream AV, as shown in figures 2 and 6 .
  • each sealing member 38a, 38b makes it possible to fill in the radial play between the bottom wall 13 and the corresponding ferrule 11, 12 of the combustion chamber 6, in order to limit the passage of air to the zones of aforementioned interface. This improves the performance of the turbomachine 1 and limits the sources of pollution.
  • each angular sector 39a, 39b can be deformed slightly so as to adapt to the actual diameter of the cylindrical part 14, 17 of the corresponding ferrule 11, 12 and of the corresponding cylindrical part 20, 21 of the bottom wall. 13. Each angular sector 39a, 39b can then optimally close off the interface zone between the bottom wall 13 and the corresponding ferrule 11, 12.
  • the figure 8 shows a second embodiment, which differs from that described with reference to figures 1 to 7 , in that the radially inner surface 45 of the outer shell 11 has a recess annular 46, the downstream axial end of which forms an annular radial shoulder 47.
  • the first sealing member 38a is housed, at least in part, in the recess 46, the downstream end of each angular sector 39a, 39b of the first sealing member 38a being able to bear on the shoulder 47.
  • each sector 39a, 39b of the first sealing member 38a and the aforementioned shoulder 47 form a baffle making it possible to limit the passage of air between these elements.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Gasket Seals (AREA)

Claims (8)

  1. Brennkammer (6) für eine Turbomaschine (1), insbesondere für das Turbostrahltriebwerk oder Turboprop-Triebwerk eines Flugzeugs, die folgendes umfasst:
    - einen radial äußeren, ringförmigen Mantel (11),
    - einen radial inneren ringförmigen Mantel (12) koaxial zum radial äußeren Mantel (11),
    - eine Bodenwand (13), die den radial äußeren Mantel (11) und den radial inneren Mantel (12) miteinander verbindet,
    - ein erstes ringförmiges Dichtungselement (38a) koaxial zu den radial inneren (12) und äußeren (11) Mantelringen, wobei das erste Dichtungselement (38a) radial zwischen der Bodenwand (13) und dem radial äußeren Mantel (11) angeordnet ist, und jedes Dichtungselement (38a, 38b) sektorisiert ist und mindestens zwei Winkelsektoren (39a, 39b) umfasst,
    - ein Wärmeschutzelement (41), das stromabwärts (AV) der Bodenwand (13) angeordnet ist, dadurch gekennzeichnet, dass das Schutzelement (41) ein Blech mit einem sich radial erstreckenden ringförmigen Teil (42) ist, dessen innere und äußere Umfangskanten durch ringförmige Ränder (43, 44) verlängert sind, die sich axial stromaufwärts (AM) erstrecken, wobei die radial äußeren (43) und inneren (44) Ränder des Schutzelements (41) radial zwischen dem äußeren Mantel (11) und der Bodenwand (6) bzw. zwischen dem inneren Mantel (12) und der Bodenwand (13) angeordnet sind.
  2. Verbrennungskammer (6) nach Anspruch 1, die ein zweites ringförmiges Dichtungselement (38b) koaxial zu den radial inneren (12) und äußeren (11) Mantelringen umfasst, wobei das zweite Dichtungselement (38b) radial zwischen der Bodenwand (13) und dem radial inneren Mantel (12) angeordnet ist.
  3. Brennkammer (6) nach Anspruch 2, in der für jedes Dichtungselement (38a, 38b) die Winkelsektoren (39a, 39b) über den Umfang verteilt sind, wobei zwischen ihnen ein Gesamtwinkelspiel zwischen 0 und 1° oder 0 und 5 mm besteht.
  4. Brennkammer (6) nach einem der Ansprüche 1 bis 3, in welcher der äußere Mantelring (11) der Brennkammer (6) einen zylindrischen Teil (14) umfasst, der einen radial äußeren zylindrischen Teil (20) der Bodenwand (13) umgibt, wobei die Bodenwand (13) des Weiteren mindestens einen radial inneren zylindrischen Teil (21) umfasst, der einen zylindrischen Teil (17) des inneren Mantels (12) der Brennkammer (6) umgibt, wobei das erste Dichtungselement (38a) zwischen dem zylindrischen Teil (14) des äußeren Mantels (11) und dem äußeren zylindrischen Teil (20) der Bodenwand (13) angeordnet ist, wobei das zweite Dichtungselement (38b) zwischen dem zylindrischen Teil (17) des inneren Mantels (12) und dem inneren zylindrischen Teil (21) der Bodenwand (13) angeordnet ist.
  5. Brennkammer (6) nach einem der Ansprüche 1 bis 4, in der jedes Dichtungselement (38a, 38b) aus einer Legierung auf Nickelbasis oder einer Legierung auf Kobaltbasis besteht.
  6. Brennkammer (6) nach einem der Ansprüche 1 bis 5, in der jedes Dichtungselement (38a, 38b) eine Dicke zwischen 0,8 und 3 mm aufweist.
  7. Brennkammer (6) nach einem der Ansprüche 1 bis 6, dadurch gekennzeichnet, dass die radial innere Oberfläche (45) des äußeren Mantelrings (11) eine ringförmige Vertiefung (46) aufweist, deren stromabwärts gelegenes axiales Ende eine ringförmige radiale Schulter (47) bildet, wobei das erste Dichtungselement (38a) zumindest teilweise in der Vertiefung (46) aufgenommen ist und das stromabwärtige Ende jedes Winkelsektors (39a, 39b) des ersten Dichtungselements (38a) in der Lage ist, an der Schulter (47) zur Anlage zu kommen.
  8. Turbomaschine (1), insbesondere Turbostrahltriebwerk oder Turboprop-Triebwerk eines Flugzeugs, mit einer Brennkammer (6) nach einem der Ansprüche 1 bis 7.
EP18700941.0A 2017-01-10 2018-01-04 Brennkammer für einen turbinenmotor Active EP3568638B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1750208A FR3061761B1 (fr) 2017-01-10 2017-01-10 Chambre de combustion pour turbomachine
PCT/FR2018/050021 WO2018130765A1 (fr) 2017-01-10 2018-01-04 Chambre de combustion pour turbomachine

Publications (2)

Publication Number Publication Date
EP3568638A1 EP3568638A1 (de) 2019-11-20
EP3568638B1 true EP3568638B1 (de) 2021-03-31

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Application Number Title Priority Date Filing Date
EP18700941.0A Active EP3568638B1 (de) 2017-01-10 2018-01-04 Brennkammer für einen turbinenmotor

Country Status (5)

Country Link
US (1) US11614234B2 (de)
EP (1) EP3568638B1 (de)
CN (1) CN110168284B (de)
FR (1) FR3061761B1 (de)
WO (1) WO2018130765A1 (de)

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JP2597800B2 (ja) * 1992-06-12 1997-04-09 ゼネラル・エレクトリック・カンパニイ ガスタービンエンジン用燃焼器
US7093440B2 (en) * 2003-06-11 2006-08-22 General Electric Company Floating liner combustor
US7051532B2 (en) * 2003-10-17 2006-05-30 General Electric Company Methods and apparatus for film cooling gas turbine engine combustors
EP1745245B1 (de) * 2004-05-05 2007-10-03 Alstom Technology Ltd Brennkammer für gasturbine
FR2897922B1 (fr) * 2006-02-27 2008-10-10 Snecma Sa Agencement pour une chambre de combustion de turboreacteur
FR2980554B1 (fr) * 2011-09-27 2013-09-27 Snecma Chambre annulaire de combustion d'une turbomachine
US10378775B2 (en) * 2012-03-23 2019-08-13 Pratt & Whitney Canada Corp. Combustor heat shield
WO2014052965A1 (en) * 2012-09-30 2014-04-03 United Technologies Corporation Interface heat shield for a combustor of a gas turbine engine
US10436446B2 (en) * 2013-09-11 2019-10-08 General Electric Company Spring loaded and sealed ceramic matrix composite combustor liner
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US10215418B2 (en) * 2014-10-13 2019-02-26 Ansaldo Energia Ip Uk Limited Sealing device for a gas turbine combustor
EP3078914A1 (de) * 2015-04-09 2016-10-12 Siemens Aktiengesellschaft Ringbrenner für einen gasturbinenmotor
US10801729B2 (en) * 2015-07-06 2020-10-13 General Electric Company Thermally coupled CMC combustor liner
US10197278B2 (en) * 2015-09-02 2019-02-05 General Electric Company Combustor assembly for a turbine engine
EP3252378A1 (de) * 2016-05-31 2017-12-06 Siemens Aktiengesellschaft Ringbrennkammer-anordnung einer gasturbine

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Also Published As

Publication number Publication date
CN110168284B (zh) 2021-02-23
WO2018130765A1 (fr) 2018-07-19
FR3061761B1 (fr) 2021-01-01
US11614234B2 (en) 2023-03-28
US20190360698A1 (en) 2019-11-28
CN110168284A (zh) 2019-08-23
FR3061761A1 (fr) 2018-07-13
EP3568638A1 (de) 2019-11-20

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