EP3359880B1 - Ringförmige brennkammer für einen turbinenmotor - Google Patents

Ringförmige brennkammer für einen turbinenmotor Download PDF

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Publication number
EP3359880B1
EP3359880B1 EP16787502.0A EP16787502A EP3359880B1 EP 3359880 B1 EP3359880 B1 EP 3359880B1 EP 16787502 A EP16787502 A EP 16787502A EP 3359880 B1 EP3359880 B1 EP 3359880B1
Authority
EP
European Patent Office
Prior art keywords
holes
injector
combustion chamber
wall
injectors
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.)
Active
Application number
EP16787502.0A
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English (en)
French (fr)
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EP3359880A1 (de
Inventor
Stéphane PASCAUD
Guillaume TALIERCIO
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 Helicopter Engines SAS
Original Assignee
Safran Helicopter Engines SAS
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Filing date
Publication date
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Priority to PL16787502T priority Critical patent/PL3359880T3/pl
Publication of EP3359880A1 publication Critical patent/EP3359880A1/de
Application granted granted Critical
Publication of EP3359880B1 publication Critical patent/EP3359880B1/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
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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
    • F23R3/26—Controlling the air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • F23R3/54—Reverse-flow combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/35—Combustors or associated equipment

Definitions

  • the present disclosure relates to an annular combustion chamber, and more particularly to an annular combustion chamber comprising a first annular wall and a second annular wall coaxial about an axis, a chamber base connecting the first and second walls, and a plurality of injectors, the first wall having first air supply holes downstream of the injectors.
  • a combustion chamber may be a turbomachine combustion chamber.
  • Combustion chambers of this type are known, for example documents FR 2 941 287 , US 2010/162712 and US 2012/304647 , or the document EP 0 569 300 which describes a guillotine device for modifying the axial position of a closure member for primary supply holes.
  • FR 2 941 287 shows a combustion chamber according to the preamble of claim 1.
  • At least one of said injectors at least three, preferably at least four, first holes having in common the first injector as the nearest injector are located equidistant from the first injector.
  • first holes are located at the same distance from the injector that is closest to them, namely the first injector, said first holes are arranged on a circular arc centered on the first injector.
  • the arc of a circle is a curve which is strictly convex in radial view and, in this case, whose concavity is turned towards the first injector.
  • the curve defines, in radial projection, a strictly convex surface, the injector being located in said surface.
  • a convex surface is such that, whatever two points belonging to the convex surface, the line segment connecting these two points belongs wholly to said convex surface.
  • a surface is strictly convex if the curve which delimits it contains no straight portion.
  • the present disclosure proposes to dimension the position of the first holes relative to the injector closest to said holes. . This results in better control of the flow and the temperature field in the combustion chamber.
  • the first injector and preferably all injectors, can be placed on the chamber bottom or on one of the annular walls, in particular on the first wall.
  • the axis of the combustion chamber is called the axis of symmetry (or quasi-symmetry) of the combustion chamber.
  • the axial direction corresponds to the direction of the axis of the combustion chamber and a radial direction is a direction perpendicular to the axis of the combustion chamber and intersecting this axis.
  • an axial plane is a plane containing the axis of the combustion chamber and a radial plane is a plane perpendicular to this axis.
  • a circumference is understood as a circle belonging to a radial plane and whose center belongs to the axis of the combustion chamber.
  • a tangential or circumferential direction is a direction tangent to a circumference; it is perpendicular to the axis of the combustion chamber but does not pass through the axis.
  • the first holes may be primary holes, i.e. holes configured to introduce fresh air, for example from the compressor, so as to delimit by turbulence, between the injectors and said holes, an anchoring zone of the flame to ensure its stability and good combustion. This zone is called primary zone.
  • the first holes may be dilution holes, i.e. holes configured to introduce fresh air, e.g. from the compressor, into the heart of the combustion chamber, distance downstream of the flame of the injector.
  • all the first holes having in common the same nearest injector are located equidistant from this injector.
  • the first holes are positioned according to their distance from the nearest injector, which allows to control the recirculation zones and therefore the temperature field in the combustion chamber. .
  • the second wall has second air supply holes downstream of the injectors.
  • the second holes may be positioned similarly to the first holes, or in a different manner.
  • the second holes preferably all the second holes, having in common the same nearest injector are located equidistant from this injector.
  • the position of the second holes can also be controlled not with respect to the chamber bottom, but with respect to the injectors which are respectively closest to them.
  • the position of the second holes can be controlled independently of the position of the first holes.
  • the first holes and the second holes having in common the same nearest injector are located equidistant from this injector. This makes it possible to have a radially homogeneous temperature field.
  • all the first and / or second holes are respectively located at the same distance from the injector that is closest to them. Thanks to these provisions, the temperature field is symmetrical of revolution around the axis of the combustion chamber. It is therefore more stable and easier to control.
  • the first holes and / or the second holes are arranged in arcs of circles centered on the injectors respectively closest.
  • the first wall is a radially outer wall and the second wall is a radially inner wall.
  • the opposite is also possible.
  • the present disclosure also relates to a turbomachine comprising an annular combustion chamber as previously described.
  • the figure 1 shows a longitudinal section of a sector of a combustion chamber 10 of an aircraft turbine engine.
  • the combustion chamber 10 is annular with a longitudinal axis X. It is delimited by a first substantially annular wall 12 around the axis X, here a radially outer wall, by a second substantially annular wall 13 about the axis X, here a radially inner wall, and a chamber bottom 14 which connects an end of the first wall 12 and the end facing the second wall 13 so as to close the upstream end of the combustion chamber 10.
  • the bottom chamber 14 is here annular.
  • the annular combustion chamber 10 further comprises a plurality of fuel injectors 16 which inject the fuel into the combustion chamber 10.
  • the injectors 16 are distributed around the longitudinal axis X. In the present embodiment, the injectors are arranged to through the chamber bottom 14. Each injector 16 defines an injection direction I.
  • first wall 12 comprises first holes formed by the first primary holes 18 and possibly the first dilution holes 20.
  • second wall 13 comprises second holes formed by the second primary holes 19 and possibly the second dilution holes 21.
  • At least three, preferably at least four of the first holes 18 having in common the first injector 16 as the closest injector are located along a strictly convex curve C, in radial view, the concavity of the curve C being turned towards the first injector 16.
  • These first holes 18 may be consecutive holes, adjacent to each other, in particular in a circumferential direction. More precisely, there are shown here six first holes 18, which are in this example first primary holes 18, arranged along the strictly convex curve C.
  • the curve C is an arc of a circle, in this case centered on the first injector 16 and typically on the injection point A of said first injector 16.
  • the first holes 18 having in common the same injector the closer, namely the first injector 16 are located equidistant D of this injector.
  • the flame 24 is stabilized by the recirculation zone 26 and fed by the suspended fuel 22.
  • the injection direction I is coplanar with the axis X of the combustion chamber.
  • FIGS. 3 and 4 are views respectively similar to those of Figures 1 and 2 in a second embodiment.
  • the elements corresponding or identical to those of the first embodiment will receive the same reference sign, to the number of hundreds, and will not be described again.
  • the injectors 116 are not arranged on the chamber bottom 114.
  • the injectors 116 are disposed on the first wall 112.
  • the injectors 116 are in position. further downstream of the chamber bottom 114.
  • the injection direction I is non-coplanar with the X axis of the combustion chamber.
  • the injection direction I has a non-zero component in the circumferential direction around the axis X.
  • the injection direction I has an axial component in the direction of the chamber bottom 114.
  • the flow of air is represented by arrows.
  • the air comes from an outlet 130 of a compressor and enters the combustion chamber 110 through the injectors 116, the first primary holes 118, the second primary holes 119, the first dilution holes 120 and the second dilution holes 121.
  • the combustion gases are discharged to the inlet 132 of a turbine.
  • At least three, preferably at least four of the first holes 118 having in common the first injector 116 as the nearest injector are located along a strictly convex curve C, in radial view, the concavity of the curve C being turned towards the first injector 116. More specifically, there are shown here six first holes 118, which are in this example first primary holes 118, four of which are arranged along the curve C strictly convex. The first three holes 118 shown in the lower right corner of the figure 4 being aligned, they can not be all three on the same strictly convex curve.
  • the location of the air inlets (here the first holes 118) is optimized, the first holes being positioned in coherence with the physical phenomena within the combustion chamber 110.
  • the flame 124 is stabilized by the recirculation zone 126 and supplied with the suspension fuel 122.
  • the position of the primary holes 118 along a strictly convex curve and whose concavity is turned towards the first injector 116, such as the curve C on the figure 4 makes it possible to stop the recirculation zone 126 and ensures better recirculation of the burnt gases as well as a homogenization of the temperature field in the combustion chamber 110.
  • the described embodiments have been detailed in the case of a single injector 16, 116 and for the first holes 18, 118 of the first wall 12, 112, similar examples could describe the distribution of the second holes on the second one. wall.
  • the second holes 19, 119 having in common the same injector 16, 116 closest can be located equidistant from this injector. This distance may be different from the distance D which separates the first holes 18, 118 of the injector 16, 116, or is equal to that shown in FIG. figure 1 .
  • the distance between the first and / or second holes and their nearest injector may vary from one injector to another, or be identical for all injectors.
  • first and / or second holes and their nearest injector may vary according to the type of holes, for example be different for the first primary holes 18, 118 and the first dilution holes 20, 120 who nevertheless have the same injector 16, 116 the nearest.
  • the first embodiment could be applied to a combustion chamber of the type of the figure 3 having injectors located on one of the annular walls, and the second embodiment could be applied to a chamber of combustion of the type of figure 1 having injectors located on the chamber floor.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (10)

  1. Ringförmige Brennkammer (10, 110), die eine erste ringförmige Wand (12, 112) und eine zweite ringförmige Wand (13, 113), die koaxial um eine Achse (X) herum sind, einen Kammerboden (14, 114), der die erste und zweite Wand (12, 13, 112, 113) verbindet, und mehrere Einspritzvorrichtungen (16, 116) umfasst, wobei die erste Wand (12, 112) erste Löcher (18, 118) zum Zuführen von Luft stromabwärts der Einspritzvorrichtungen (16, 116) umfasst, wobei die Brennkammer (10, 110) dadurch gekennzeichnet ist, dass sich für zumindest eine erste der Einspritzvorrichtungen (16, 116) zumindest drei der ersten Löcher (18, 118), welche die erste Einspritzvorrichtung als nächste Einspritzvorrichtung gemeinsam haben, im gleichen Abstand (D) zu der ersten Einspritzvorrichtung (16, 116) befinden.
  2. Brennkammer (110) nach Anspruch 1, wobei die erste Einspritzvorrichtung (116) auf der ersten Wand (112) platziert ist.
  3. Brennkammer (10) nach Anspruch 1 oder 2, wobei sich zumindest vier der ersten Löcher (18), welche die erste Einspritzvorrichtung als nächste Einspritzvorrichtung gemeinsam haben, im gleichen Abstand (D) zu dieser Einspritzvorrichtung (16) befinden.
  4. Brennkammer (10) nach einem der Ansprüche 1 bis 3, wobei sich alle ersten Löcher (18), die eine gleiche Einspritzvorrichtung (16) als nächste gemeinsam haben, im gleichen Abstand (D) zu dieser Einspritzvorrichtung befinden.
  5. Brennkammer (10) nach einem der Ansprüche 1 bis 4, wobei die zweite Wand (13) zweite Löcher (19) zum Zuführen von Luft stromabwärts der Einspritzvorrichtungen (16) umfasst, und wobei sich die zweiten Löcher (19), bevorzugt alle zweiten Löcher (16), die eine gleiche Einspritzvorrichtung (16) als nächste gemeinsam haben, im gleichen Abstand (D) zu dieser Einspritzvorrichtung befinden.
  6. Brennkammer (10) nach Anspruch 5, wobei sich die ersten Löcher (18) und die zweiten Löcher (19), die eine gleiche Einspritzvorrichtung (16) als nächste gemeinsam haben, im gleichen Abstand (D) zu dieser Einspritzvorrichtung befinden.
  7. Brennkammer (10) nach einem der Ansprüche 1 bis 6, wobei sich alle ersten und/oder zweiten Löcher (18, 19) jeweils in einem gleichen Abstand zu der Einspritzvorrichtung (16) befinden, die ihnen am nächsten ist.
  8. Brennkammer (10) nach einem der Ansprüche 1 bis 7, wobei die ersten Löcher (18) und/oder die zweiten Löcher (19) gemäß zentrierten Kreisbögen auf den jeweils nächsten Einspritzvorrichtungen angeordnet sind.
  9. Brennkammer (10) nach einem der Ansprüche 1 bis 8, wobei die erste Wand (12) eine radial äußere Wand ist und die zweite Wand (14) eine radial innere Wand ist.
  10. Turbomaschine, die eine Brennkammer (10) nach einem der Ansprüche 1 bis 9 umfasst.
EP16787502.0A 2015-10-06 2016-10-04 Ringförmige brennkammer für einen turbinenmotor Active EP3359880B1 (de)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PL16787502T PL3359880T3 (pl) 2015-10-06 2016-10-04 Pierścieniowa komora spalania do maszyny wirowej

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1559491A FR3042023B1 (fr) 2015-10-06 2015-10-06 Chambre de combustion annulaire pour turbomachine
PCT/FR2016/052539 WO2017060605A1 (fr) 2015-10-06 2016-10-04 Chambre de combustion annulaire pour turbomachine

Publications (2)

Publication Number Publication Date
EP3359880A1 EP3359880A1 (de) 2018-08-15
EP3359880B1 true EP3359880B1 (de) 2019-08-21

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EP16787502.0A Active EP3359880B1 (de) 2015-10-06 2016-10-04 Ringförmige brennkammer für einen turbinenmotor

Country Status (10)

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US (1) US10895383B2 (de)
EP (1) EP3359880B1 (de)
JP (1) JP2018534518A (de)
KR (1) KR20180063120A (de)
CN (1) CN108139077B (de)
CA (1) CA3000922A1 (de)
FR (1) FR3042023B1 (de)
PL (1) PL3359880T3 (de)
RU (1) RU2718375C2 (de)
WO (1) WO2017060605A1 (de)

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Publication number Priority date Publication date Assignee Title
US3671171A (en) 1970-11-27 1972-06-20 Avco Corp Annular combustors
CA2076102C (en) 1991-09-23 2001-12-18 Stephen John Howell Aero-slinger combustor
FR2690977B1 (fr) 1992-05-06 1995-09-01 Snecma Chambre de combustion comportant des passages reglables d'admission de comburant primaire.
JP3612331B2 (ja) * 1993-06-01 2005-01-19 プラット アンド ホイットニー カナダ，インコーポレイテッド 半径方向に取り付けられる空気噴出型燃料噴射弁
FR2751054B1 (fr) 1996-07-11 1998-09-18 Snecma Chambre de combustion anti-nox a injection de carburant de type annulaire
US6155056A (en) * 1998-06-04 2000-12-05 Pratt & Whitney Canada Corp. Cooling louver for annular gas turbine engine combustion chamber
US6367262B1 (en) * 2000-09-29 2002-04-09 General Electric Company Multiple annular swirler
FR2856467B1 (fr) * 2003-06-18 2005-09-02 Snecma Moteurs Chambre de combustion annulaire de turbomachine
US7509809B2 (en) * 2005-06-10 2009-03-31 Pratt & Whitney Canada Corp. Gas turbine engine combustor with improved cooling
FR2903169B1 (fr) * 2006-06-29 2011-11-11 Snecma Dispositif d'injection d'un melange d'air et de carburant, chambre de combustion et turbomachine munies d'un tel dispositif
FR2911667B1 (fr) * 2007-01-23 2009-10-02 Snecma Sa Systeme d'injection de carburant a double injecteur.
US8171634B2 (en) * 2007-07-09 2012-05-08 Pratt & Whitney Canada Corp. Method of producing effusion holes
FR2919380B1 (fr) * 2007-07-26 2013-10-25 Snecma Chambre de combustion d'une turbomachine.
US8616004B2 (en) 2007-11-29 2013-12-31 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
FR2941287B1 (fr) 2009-01-19 2011-03-25 Snecma Paroi de chambre de combustion de turbomachine a une seule rangee annulaire d'orifices d'entree d'air primaire et de dilution
US8171740B2 (en) * 2009-02-27 2012-05-08 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
FR2948988B1 (fr) 2009-08-04 2011-12-09 Snecma Chambre de combustion de turbomachine comprenant des orifices d'entree d'air ameliores
US9080770B2 (en) * 2011-06-06 2015-07-14 Honeywell International Inc. Reverse-flow annular combustor for reduced emissions
CN204460285U (zh) 2014-12-03 2015-07-08 中国航空工业集团公司金城南京机电液压工程研究中心 一种环形回流燃烧室火焰筒

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Also Published As

Publication number Publication date
EP3359880A1 (de) 2018-08-15
PL3359880T3 (pl) 2019-12-31
KR20180063120A (ko) 2018-06-11
CA3000922A1 (fr) 2017-04-13
US20180299127A1 (en) 2018-10-18
CN108139077B (zh) 2020-09-15
CN108139077A (zh) 2018-06-08
WO2017060605A1 (fr) 2017-04-13
US10895383B2 (en) 2021-01-19
RU2018116261A3 (de) 2020-01-21
RU2718375C2 (ru) 2020-04-02
JP2018534518A (ja) 2018-11-22
FR3042023A1 (fr) 2017-04-07
FR3042023B1 (fr) 2020-06-05
RU2018116261A (ru) 2019-11-07

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