EP2003399B1 - Chambre de combustion de turbomachine à circulation hélicoïdale de l'air - Google Patents

Chambre de combustion de turbomachine à circulation hélicoïdale de l'air Download PDF

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Publication number
EP2003399B1
EP2003399B1 EP08158059.9A EP08158059A EP2003399B1 EP 2003399 B1 EP2003399 B1 EP 2003399B1 EP 08158059 A EP08158059 A EP 08158059A EP 2003399 B1 EP2003399 B1 EP 2003399B1
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EP
European Patent Office
Prior art keywords
combustion chamber
pilot
wall
air
turbomachine
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
EP08158059.9A
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German (de)
English (en)
French (fr)
Other versions
EP2003399A3 (fr
EP2003399A2 (fr
Inventor
Laurent Bernard Cameriano
Michel Pierre Cazalens
Sylvain Duval
Romain Nicolas Lunel
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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SNECMA SAS
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Publication of EP2003399A3 publication Critical patent/EP2003399A3/fr
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Publication of EP2003399B1 publication Critical patent/EP2003399B1/fr
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing

Definitions

  • the present invention relates to the general field of combustion chambers of an aeronautical or terrestrial turbomachine.
  • An aeronautical or terrestrial turbomachine is typically formed of an assembly comprising in particular an annular compression section intended to compress air passing through the turbomachine, an annular combustion section disposed at the outlet of the compression section and in which the air coming from the compression section is mixed with fuel for burning, and an annular turbine section disposed at the outlet of the combustion section and a rotor is rotated by gases from the combustion section.
  • a combustion chamber according to the state of the art is disclosed in FR 2 695 460 A1 .
  • the compression section is in the form of a plurality of stages of movable wheels each carrying blades which are arranged in an annular channel through which the air of the turbomachine and whose section decreases from upstream to downstream.
  • the combustion section includes a combustion chamber in the form of an annular channel in which the compressed air is mixed with fuel for burning.
  • the turbine section it is formed by a plurality of stages of moving wheels each carrying blades which are arranged in an annular channel through which the combustion gases pass.
  • the circulation of air through this assembly is generally carried out as follows: the compressed air from the last stage of the compression section has a natural rotational movement with an inclination of the order of 35 ° to 45 ° ° with respect to the longitudinal axis of the turbomachine, tilt which varies according to the speed of the turbomachine (speed of rotation).
  • this compressed air is straightened in the longitudinal axis of the turbomachine (that is to say that the inclination of the air with respect to the longitudinal axis of the turbomachine is brought back at 0 °) via an air rectifier.
  • the air in the combustion chamber is then mixed with fuel so as to ensure satisfactory combustion and the gases from this combustion continue a course generally along the longitudinal axis of the turbomachine to reach the turbine section.
  • the combustion gases are reoriented by a distributor to present a gyratory movement with an inclination greater than 70 ° with respect to the longitudinal axis of the turbine engine.
  • Such inclination is essential to produce the angle of attack required for the mechanical force driving in rotation of the moving wheel of the first stage of the turbine section.
  • Such angular distribution of the air passing through the turbomachine has many disadvantages. Indeed, the air that naturally leaves the last stage of the compression section with an angle between 35 ° and 45 ° is successively rectified (angle reduced to 0 °) at its entry into the combustion section and then reoriented with an angle greater than 70 ° at its entry into the turbine section. These successive angular modifications of the distribution of air through the turbomachine require intense aerodynamic forces produced by the rectifier of the compression section and the distributor of the turbine section, aerodynamic forces which are particularly detrimental to the overall efficiency of the turbomachine. the turbomachine.
  • the present invention aims to overcome the aforementioned drawbacks by providing a turbomachine combustion chamber that can be powered by an air that has a rotational movement with respect to the longitudinal axis of the turbomachine.
  • the combustion chamber according to the invention can be supplied with air having a rotational movement about the longitudinal axis of the turbomachine.
  • the natural inclination of the air at the outlet of the compression section of the turbomachine can therefore be maintained through the combustion chamber.
  • the aerodynamic force required for rotating the first stage of the turbine section of the turbomachine is considerably reduced. This sharp decrease in aerodynamic forces generates a gain in efficiency of the turbomachine.
  • the rectifier of the compression section and the distributor of the turbine section can be simplified or even eliminated, which represents a saving in weight and a reduction in production costs.
  • pilot cavities which are carburized only for the idling speeds of the turbomachine, provides a stabilization of the combustion flame for all operating conditions of the turbomachine.
  • each pilot cavity is closed at its upstream end and open at its downstream end.
  • each pilot cavity is delimited circumferentially by two substantially radial partitions, one of these partitions comprising a plurality of air injection orifices opening towards the outside of the combustion chamber and opening in said pilot cavity.
  • the other partition of each pilot cavity has, in cross section, a substantially curvilinear section.
  • the full-throttle injectors are offset axially downstream relative to the pilot injectors. Indeed, the flame from the pilot injectors needs a residence time in the combustion chamber which is higher than the flame from the injectors full throttle.
  • the combustion chamber may be devoid of wall connecting transversely upstream longitudinal ends of the inner and outer walls.
  • the absence of such a wall makes it possible to preserve as much as possible the rotation of the air coming from the compression section of the turbomachine.
  • the fuel injection systems are devoid of associated air systems.
  • the combustion chamber may further comprise an inner annular fairing which is mounted on the inner wall in the upstream extension thereof and an outer annular fairing which is mounted on the outer wall in the upstream extension thereof.
  • the invention also relates to a turbomachine comprising a combustion chamber as defined above.
  • the turbomachine partially shown on the figure 1 has a longitudinal axis XX. According to this axis, it comprises in particular: an annular compression section 100, an annular combustion section 200 disposed at the outlet of the section of compression 100 according to the direction of flow of the air passing through the turbomachine, and an annular turbine section 300 disposed at the outlet of the combustion section 200.
  • the air injected into the turbomachine therefore passes successively through the compression section 100, then the combustion section 200 and finally the turbine section 300.
  • the compression section 100 is in the form of a plurality of stages of movable wheels 102 each carrying blades 104 (only the last stage of the compression section is shown in FIG. figure 1 ).
  • the blades 104 of these stages are disposed in an annular channel 106 through which air flows through the turbomachine and whose section decreases from upstream to downstream. Thus, as the air injected into the turbomachine passes through the compression section, it is more and more compressed.
  • the combustion section 200 is also in the form of an annular channel in which the compressed air from the compression section 100 is mixed with fuel for burning there.
  • the combustion section comprises a combustion chamber 202 inside which is burned the air / fuel mixture (this chamber is detailed later).
  • the combustion section 200 also comprises a turbomachine casing formed of an outer annular casing 204 centered on the longitudinal axis XX of the turbomachine and an inner annular casing 206 which is fixed coaxially inside the casing. outer envelope. An annular space 208 formed between these two envelopes 204, 206 receives compressed air from the compression section 100 of the turbomachine.
  • the turbomachine section 300 of the turbomachine is formed by a plurality of stages of movable wheels 302 each carrying blades 304 (only the first stage of the turbine section is shown in FIG. figure 1 ).
  • the blades 304 of these stages are arranged in an annular channel 306 traversed by the gases coming from the combustion section 200.
  • the gases coming from the combustion section must have an inclination relative to the longitudinal axis XX of the turbomachine which is sufficient to rotate the different stages of the turbine section. turbine.
  • a distributor 308 is mounted directly downstream of the combustion chamber 202 and upstream of the first stage 302 of the turbine section 300.
  • This distributor 308 consists of a plurality of fixed radial vanes 310 of which inclination with respect to the longitudinal axis XX of the turbomachine makes it possible to give the gases coming from the combustion section 200 the inclination necessary for driving in rotation the different stages of the turbine section.
  • the distribution of the air successively passing through the compression section 100, the combustion section 200 and the turbine section 300 takes place as follows.
  • the compressed air from the last stage 102 of the compression section 100 naturally has a gyratory movement with an inclination of the order of 35 ° to 45 ° relative to the longitudinal axis X-X of the turbomachine.
  • this inclination angle is reduced to 0 °.
  • the gases resulting from the combustion are redirected by the blades 310 of the distributor 308 of the latter to give them a gyratory movement with an inclination with respect to the longitudinal axis XX which is greater than 70 °.
  • a new architecture of the combustion chamber 202 which can be powered by an air having a rotational movement about the longitudinal axis X-X of the turbomachine.
  • an architecture it is possible to maintain the natural inclination of the compressed air from the last stage of the compression section without having to straighten it in the longitudinal axis X-X.
  • the stationary blades 310 of the distributor 308 of the turbine section 300 it is no longer necessary for the stationary blades 310 of the distributor 308 of the turbine section 300 to have such a large inclination to produce the angle of attack required for the mechanical driving force in rotation of the moving wheel. 302 of the first stage of the turbine section.
  • the combustion chamber 202 comprises an inner annular wall 212 centered on the longitudinal axis XX of the turbomachine, and an outer annular wall 214 also centered on the longitudinal axis XX and surrounding the inner wall of the engine. to define with it an annular space 216 forming a combustion chamber.
  • the combustion chamber 202 further comprises at least one air inlet opening 218 which opens into the combustion chamber 216 at the upstream end thereof and in a substantially longitudinal direction.
  • the section of this air intake opening is adapted to ensure the operation of the combustion chamber.
  • this air inlet opening 218 is formed between the upstream ends of the inner and outer walls 212 and 214 of the combustion chamber.
  • the combustion chamber 202 also comprises a plurality of fuel injection systems 220 distributed on the outer wall 214 around the longitudinal axis XX of the turbomachine and opening into the combustion chamber 216 in a substantially radial direction .
  • the fuel injection systems 220 comprise pilot injectors 220a circumferentially alternating with full-throttle injectors 220b, the full-throttle injectors preferably being offset axially downstream relative to the pilot injectors.
  • the pilot injectors 220a provide ignition and idle phases of the turbomachine and the 220b full-throttle injectors are involved in the take-off, climb and cruise phases.
  • the pilot injectors are fueled continuously while the full-throttle injectors are only fed beyond a certain determined speed.
  • the fuel injection systems 220 are devoid of associated air systems such as air swirlers which make it possible, in a manner known per se, to generate a rotary air flow. inside the combustion chamber in order to stabilize the combustion flame.
  • pilot and full throttle injectors of the combustion chamber are of very simple design and very reliable operation since they are reduced to their simplest function, namely to inject fuel.
  • the pilot injectors 220a are of the same type as the full throttle injectors 220b.
  • the outer wall 214 of the combustion chamber comprises a plurality of pilot cavities 222 which are regularly distributed around the longitudinal axis X-X.
  • each pilot cavity 222 extends, firstly longitudinally between the two longitudinal ends (upstream and downstream) of the outer wall 214, and secondly radially outwardly thereof.
  • the outer wall 214 is profiled with a plurality of cavities 222 protruding outwardly from the wall.
  • pilot cavities 222 are each delimited circumferentially by two partitions 224 which each project radially outwardly with respect to the outer wall 214. As shown in FIGS. figures 2 and 5 , one of these partitions has a plurality of air injection orifices 226 which make it possible to inject air outside the combustion chamber into the pilot cavity in a circumferential direction.
  • the circumferential injection of air is performed in the same direction of rotation (that of the needles of a watch for the exemplary embodiment of figures 2 and 3 ) for all the pilot cavities 222 of the combustion chamber. Moreover, the direction of rotation for the circumferential injection of air into these pilot cavities is that of the compressed air coming from the compression section of the turbomachine.
  • pilot cavities 222 are supplied with fuel via pilot injectors 220a, each of which opens radially into one of these cavities. As for the full-throttle injectors 220b, they each open radially into the combustion chamber between two adjacent pilot cavities.
  • Each pilot cavity 222 is preferably closed at its upstream end by a radial partition 228 and open at its downstream end (see especially the figures 2 and 5 ). Thus, the air entering the combustion chamber 216 through its air inlet opening 218 does not disturb the flow of air introduced into the pilot cavities 222 by the air injection orifices 226.
  • the operation of the combustion chamber is as follows: the compressed air coming from the compression section 100 and which is rotated about the longitudinal axis XX enters the combustion section 200. This air is divided into two flows: “internal” flow and "external” flow.
  • the external flow bypasses the combustion chamber 202 and feeds the pilot cavities 222 after cooling the outer wall 214 of the combustion chamber and the outer casing 204 of the combustion section. This external air is injected into these pilot cavities via the air injection orifices 226 in the direction of rotation of the air at its entry into the combustion section. In these pilot cavities, the air is mixed and burnt with the fuel injected by the pilot injectors 220a.
  • the internal flow that represents the main flow it enters the combustion chamber 216 through the air inlet opening 218 to be mixed and burned fuel injected by the full-throttle injectors 220b. Stabilization of the combustion flame is obtained thanks to the "carburation" of the pilot cavities.
  • each pilot cavity 222 which is not provided with air injection orifices has, in cross-section, a substantially curvilinear section (unlike the other wall which is substantially flat).
  • the curvature of these walls makes it possible to accompany the rotational movement of the air injected into the pilot cavities by the air injection orifices 226.
  • the two longitudinal partitions 224 circumferentially delimiting each pilot cavity 222 are substantially flat and each extend in a radial direction.
  • the number and the geometric dimensions of the pilot cavities 222 of the combustion chamber may vary according to the needs. The same is true of the number, the dimensions and the positioning of the air injection orifices 226 in these cavities.
  • the combustion chamber 202 may also comprise an internal annular fairing 230 which is mounted on the inner wall 212 in the upstream extension thereof and an outer annular fairing 232 which is mounted on the outer wall 214 in the upstream extension thereof.
  • the presence of these shrouds 230, 232 makes it possible to regulate the flow rate of air entering the combustion chamber 202 and that bypassing it.
  • the outer wall 214 of the combustion chamber may comprise at its downstream end an annular flange 234 extending radially outwardly of the wall, this flange being provided with a plurality of holes 236 regularly distributed around the longitudinal axis XX and intended to supply cooling air to the turbine section 300.

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)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Fuel-Injection Apparatus (AREA)
  • Supercharger (AREA)
EP08158059.9A 2007-06-14 2008-06-11 Chambre de combustion de turbomachine à circulation hélicoïdale de l'air Active EP2003399B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0755761A FR2917487B1 (fr) 2007-06-14 2007-06-14 Chambre de combustion de turbomachine a circulation helicoidale de l'air

Publications (3)

Publication Number Publication Date
EP2003399A2 EP2003399A2 (fr) 2008-12-17
EP2003399A3 EP2003399A3 (fr) 2013-07-31
EP2003399B1 true EP2003399B1 (fr) 2014-04-30

Family

ID=39004879

Family Applications (1)

Application Number Title Priority Date Filing Date
EP08158059.9A Active EP2003399B1 (fr) 2007-06-14 2008-06-11 Chambre de combustion de turbomachine à circulation hélicoïdale de l'air

Country Status (8)

Country Link
US (1) US7673456B2 (he)
EP (1) EP2003399B1 (he)
JP (1) JP5084626B2 (he)
CN (1) CN101324344B (he)
CA (1) CA2634615C (he)
FR (1) FR2917487B1 (he)
IL (1) IL192052A (he)
RU (1) RU2478880C2 (he)

Families Citing this family (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6110854B2 (ja) * 2011-08-22 2017-04-05 トクァン，マジェドＴＯＱＡＮ， Ｍａｊｅｄ ガス・タービン・エンジンで使用するための予混合燃料空気を用いた接線方向環状燃焼器
CN103470376A (zh) * 2013-09-23 2013-12-25 蔡肃民 红外线发生器
RU182644U1 (ru) * 2018-03-28 2018-08-24 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения им. П.И. Баранова" Кольцевая камера сгорания малоразмерного газотурбинного двигателя
US11378277B2 (en) 2018-04-06 2022-07-05 General Electric Company Gas turbine engine and combustor having air inlets and pilot burner
FR3081494B1 (fr) 2018-05-28 2020-12-25 Safran Aircraft Engines Module de combustion de turbomachine a gaz avec butee de fond de chambre
US11181269B2 (en) * 2018-11-15 2021-11-23 General Electric Company Involute trapped vortex combustor assembly
CN112577069B (zh) * 2020-12-17 2022-03-29 中国科学院工程热物理研究所 一种适用于小头部倾斜角下的斜流燃烧室侧壁面结构
CN113154456B (zh) * 2021-04-15 2022-06-21 中国航发湖南动力机械研究所 回流燃烧室机匣头部结构及其制造方法和发动机燃烧室
CN113739207B (zh) * 2021-09-22 2022-04-29 西北工业大学 一种采用气动内柱的旋转爆震燃烧室
CN113803744B (zh) * 2021-09-27 2023-03-10 中国联合重型燃气轮机技术有限公司 燃烧室入料装置及入料系统

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CA2076102C (en) * 1991-09-23 2001-12-18 Stephen John Howell Aero-slinger combustor
FR2695460B1 (fr) * 1992-09-09 1994-10-21 Snecma Chambre de combustion de turbomachine à plusieurs injecteurs.
US5791148A (en) * 1995-06-07 1998-08-11 General Electric Company Liner of a gas turbine engine combustor having trapped vortex cavity
DE19549143A1 (de) * 1995-12-29 1997-07-03 Abb Research Ltd Gasturbinenringbrennkammer
JPH09222228A (ja) * 1996-02-16 1997-08-26 Toshiba Corp ガスタービン燃焼器
JP3673009B2 (ja) * 1996-03-28 2005-07-20 株式会社東芝 ガスタービン燃焼器
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FR2920523B1 (fr) * 2007-09-05 2009-12-18 Snecma Chambre de combustion de turbomachine a circulation helicoidale de l'air.

Also Published As

Publication number Publication date
CN101324344A (zh) 2008-12-17
IL192052A0 (en) 2009-02-11
US20080307792A1 (en) 2008-12-18
EP2003399A3 (fr) 2013-07-31
JP2008309466A (ja) 2008-12-25
CN101324344B (zh) 2011-08-17
FR2917487B1 (fr) 2009-10-02
FR2917487A1 (fr) 2008-12-19
US7673456B2 (en) 2010-03-09
EP2003399A2 (fr) 2008-12-17
CA2634615A1 (fr) 2008-12-14
RU2008124152A (ru) 2009-12-20
JP5084626B2 (ja) 2012-11-28
RU2478880C2 (ru) 2013-04-10
CA2634615C (fr) 2014-08-05
IL192052A (he) 2011-07-31

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