EP2034245B1 - Gasturbinenbrennkammer mit schraubenförmigem Luftumlauf - Google Patents

Gasturbinenbrennkammer mit schraubenförmigem Luftumlauf Download PDF

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Publication number
EP2034245B1
EP2034245B1 EP08163522A EP08163522A EP2034245B1 EP 2034245 B1 EP2034245 B1 EP 2034245B1 EP 08163522 A EP08163522 A EP 08163522A EP 08163522 A EP08163522 A EP 08163522A EP 2034245 B1 EP2034245 B1 EP 2034245B1
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EP
European Patent Office
Prior art keywords
combustion chamber
wall
air
turbomachine
longitudinal axis
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Active
Application number
EP08163522A
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English (en)
French (fr)
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EP2034245A1 (de
Inventor
Michel Pierre Cazalens
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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    • 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
    • F23R3/58—Cyclone or vortex type combustion chambers
    • 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.
  • 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 a satisfactory combustion and the gases resulting 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 ° relative to the longitudinal axis of the turbomachine.
  • 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.
  • a turbomachine combustion chamber according to the preamble of claim 1 is known from the document U.S. 5,025,622 A .
  • 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 is supplied with air via the internal and external cavities in a substantially circumferential direction.
  • the combustion chamber according to the invention can thus 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 design of the high-pressure turbine distributor can be simplified and the aerodynamic force required to bring the flow in the axis of the turbomachine substantially decreased. 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 being simplified, this can lead to a saving in weight and a reduction in production costs.
  • certain internal and external steps comprise a substantially radial wall provided with a plurality of air injection orifices opening towards the outside of the combustion chamber and opening into the adjacent internal or external cavity.
  • the internal and external steps comprise another wall which has, in cross section, a substantially curvilinear section.
  • the fuel injection systems comprise pilot injectors alternating circumferentially with full-throttle injectors.
  • the full-throttle injectors are preferably axially offset downstream relative to the pilot injectors.
  • 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 fuel injection systems are devoid of associated air systems (which generally allow the air to be rotated so as to create a recirculation in order to stabilize the combustion flame ).
  • 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. Along this axis, it comprises in particular an annular compression section 100, an annular combustion section 200 disposed at the outlet of the compression section 100 in the direction of flow of the air passing through the turbomachine, and a section annular turbine 300 disposed at the output of the combustion section 200.
  • the air injected into the turbomachine thus successively passes 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 whose inclination relative 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 rotating 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, an outer annular wall 214 also centered on the longitudinal axis XX and surrounding the inner wall of to define therewith an annular space 216 forming a combustion focus, and a transverse annular wall 218 (called chamber bottom) transversely connecting the longitudinal ends upstream of the inner and outer walls.
  • the internal wall 212 of the combustion chamber comprises a plurality of internal steps (or steps) 220 which are regularly distributed around the longitudinal axis X-X. Each of these internal steps 220 extends, firstly longitudinally between the two longitudinal ends (upstream and downstream) of the inner wall, and secondly radially outwardly thereof.
  • the inner surface of the inner wall 212 is profiled with a plurality of steps 220 protruding outwardly from the wall. Furthermore, internal cavity 222 designates the circumferential spacing that is defined between two adjacent internal steps 220.
  • the outer wall 214 of the combustion chamber comprises a plurality of steps (or steps) external 224 evenly distributed around the longitudinal axis X-X.
  • Each external step 224 extends, firstly longitudinally between the two longitudinal ends of the outer wall, and secondly radially inwardly thereof.
  • the outer surface of the outer wall 214 is profiled with a plurality of steps 224 projecting inwardly from the wall.
  • External cavity 226 denotes the circumferential spacing that is defined between two adjacent external steps 224.
  • some of the internal cavities 222 and some of the external cavities 226 are supplied with fuel in a substantially radial direction.
  • the combustion chamber 202 also comprises a plurality of fuel injection systems 228 distributed on the inner walls 212 and outer 214 around the longitudinal axis XX of the turbomachine and opening into the combustion chamber. combustion 216 in a substantially radial direction.
  • the fuel injection systems 228 radially open into some of the internal cavities 222 and some of the outer cavities 226.
  • the fuel injection systems 228 open into all the external cavities 226 and into only one internal cavity 222 out of two.
  • all the internal cavities and all the external cavities can be supplied with fuel; only one external cavity out of two and all internal cavities are fueled; etc.
  • the principle governing the choice of the supply configuration of these cavities is to optimize the performance of the combustion chamber for each point of the flight envelope.
  • fuel injection systems 228 comprise pilot injectors 228a alternating circumferentially with full-throttle injectors 228b.
  • the fuel injection systems 228 supplying the external cavities 226 do indeed comprise an alternation of pilot injectors 228a with full-throttle injectors, and the fuel injection systems 228 supplying the internal cavities 222 comprise full-throttle injectors and injectors. pilot injectors.
  • the pilot injectors 228a provide ignition and idle phases of the turbomachine and the full-throttle injectors 228b are involved in the take-off, climb and cruise phases.
  • the pilot injectors are fueled continuously while the takeoff injectors are only fed beyond a certain determined regime.
  • the fuel injection systems 228 do not have 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.
  • pilot injectors 228a are of the same type as the full-throttle injectors 228b.
  • the full-throttle injectors 228b can be axially offset downstream relative to the pilot injectors 228a.
  • At least some of the internal cavities 222 and some of the external cavities 226 are supplied with air outside the combustion chamber 202 in a same substantially circumferential direction.
  • the internal cavities 222 and external 226 which are supplied with air by means of a plurality of air injection orifices 230 formed in a substantially radial wall 232 of the internal 220 and external steps 224 corresponding.
  • These air injection orifices 230 open towards the outside of the combustion chamber 202 and open into the corresponding internal or external cavity in a substantially circumferential direction.
  • the circumferential injection of air into the combustion chamber 216 is made in the same direction of rotation (that of the needles of a watch for the embodiment of figures 2 and 3 ) for all internal cavities 222 and external 226 of the combustion chamber. Moreover, the direction of rotation for the circumferential injection of air into these cavities is that of the compressed air coming from the compression section of the turbomachine.
  • the air supply of the combustion chamber 206 is only achieved by means of the air injection orifices 230 opening into some of the internal and external cavities in a circumferential direction (a very small proportion of air also enters the combustion chamber through multiperforation holes made in the walls 212, 214 and 218 of the combustion chamber for cooling these walls, these holes are not shown in the figures).
  • the internal and external cavities that are supplied with fuel are not necessarily homogeneous with respect to their radial dimension (that is to say the height of the corresponding step) and circumferential so as to be able to vary the residence time following the cavity considered.
  • the height of the steps is not necessarily constant over the entire length of the wall (that is to say between its upstream and downstream ends).
  • the air flow supplying these cavities may vary depending on the cavity considered.
  • the operation of the combustion chamber is as follows: the compressed air coming from the compression section 100 and rotating about the longitudinal axis XX enters the combustion section 200. This air bypasses the combustion chamber 202 and supplies at least some of the internal and external cavities 222 and 226 after cooling the walls and casings of the combustion chamber. This air is injected into these cavities via the air injection orifices 230 in the direction of rotation of the air at its entry into the combustion section. In some of these air-fed cavities, the air is mixed and burned with the fuel injected by the fuel injection systems 228.
  • the internal and external steps 224 224 of the combustion chamber comprise another wall 232 '(opposite to that 232 provided with air injection orifices) which extends in a substantially circumferential direction and which has, in section transverse, a substantially curvilinear section (unlike the wall 232 which is substantially flat and radial).
  • the curvature of this wall makes it possible to form a ramp to accompany the rotational movement of the air injected into the cavities via the air injection orifices 230.
  • any other wall form (planar or curvilinear) can be envisaged. .
  • the number and the geometrical dimensions of the internal and external cavities 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 in these cavities, as well as the relative circumferential position of the fuel injection systems with respect to the internal and external steps.
  • the inner wall 212 and the outer wall 214 of the combustion chamber may each comprise at their downstream end an annular flange, respectively 234 and 236, which is provided with a plurality of holes 238 regularly distributed about the longitudinal axis XX and for supplying cooling air to the turbine section 300.

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  • 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)

Claims (7)

  1. Brennkammer (202) einer Turbomaschine, die folgendes umfaßt:
    eine ringförmige Innenwand (212) mit der Längsachse (X-X),
    eine ringförmige Außenwand (214), die um die Längsachse zentriert ist und die Innenwand derart umgibt, daß sie mit dieser einen einen Brennraum bildenden ringförmigen Raum (216) begrenzt,
    eine ringförmige Querwand (218), welche die stromaufwärtigen Längsenden der Innen- und Außenwände quer verbindet, sowie
    eine Vielzahl von Treibstoffeinspritzsystemen (228),
    wobei die Innenwand (212) eine Vielzahl von Innenstufen (220) umfaßt, die um die Längsachse gleichmäßig verteilt sind, wobei jede Innenstufe sich in Längsrichtung zwischen den beiden Längsenden der Innenwand und radial zu deren Außenseite erstreckt, wobei der Abstand in Umfangsrichtung zwischen zwei benachbarten Innenstufen einen Innenhohlraum (222) definiert,
    wobei die Außenwand (214) eine Vielzahl von Außenstufen (224) umfaßt, die um die Längsachse gleichmäßig verteilt sind, wobei jede Außenstufe sich in Längsrichtung zwischen den beiden Längsenden der Außenwand und radial zu deren Innenseite erstreckt, wobei der Abstand in Umfangsrichtung zwischen zwei benachbarten Außenstufen einen Außenhohlraum (226) definiert, und
    von denen einige Innen- und Außenhohlräume mit außerhalb der Brennkammer befindlicher Luft in einer gleichen Richtung, im wesentlichen Umfangsrichtung beaufschlagt werden,
    dadurch gekennzeichnet, daß diese einigen Innen- und Außenhohlräume auch mit Treibstoff in einer im wesentlichen radialen Richtung beaufschlagt werden.
  2. Brennkammer nach Anspruch 1, wobei einige Innenstufen (220) und Außenstufen (224) eine im wesentlichen radiale Wand (232) aufweisen, die mit einer Vielzahl von Lufteinspritzöffnungen (230) versehen ist, welche sich zur Außenseite der Brennkammer öffnen und in den benachbarten Innen- oder Außenhohlraum münden.
  3. Brennkammer nach Anspruch 2, wobei die Innenstufen (220) und Außenstufen (224) eine weitere Wand (232') aufweisen, die im Querschnitt einen im wesentlichen bogenförmigen Querschnitt aufweist.
  4. Brennkammer nach einem der Ansprüche 1 bis 3, wobei die Treibstoffeinspritzsysteme (228) Pilotinjektoren (228a) umfassen, die sich in Umfangsrichtung mit Vollgasinjektoren (228b) abwechseln.
  5. Brennkammer nach Anspruch 4, wobei die Vollgasinjektoren (228b) in stromabwärtiger Richtung gegenüber den Pilotinjektoren (228a) axial versetzt sind.
  6. Brennkammer nach einem der Ansprüche 1 bis 5, wobei die Treibstoffeinspritzsysteme (228) keine zugeordneten Luftsysteme haben.
  7. Turbomaschine, dadurch gekennzeichnet, daß sie eine Brennkammer (202) nach einem der Ansprüche 1 bis 6 umfaßt.
EP08163522A 2007-09-05 2008-09-02 Gasturbinenbrennkammer mit schraubenförmigem Luftumlauf Active EP2034245B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0757356A FR2920523B1 (fr) 2007-09-05 2007-09-05 Chambre de combustion de turbomachine a circulation helicoidale de l'air.

Publications (2)

Publication Number Publication Date
EP2034245A1 EP2034245A1 (de) 2009-03-11
EP2034245B1 true EP2034245B1 (de) 2010-04-21

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US (1) US7614234B2 (de)
EP (1) EP2034245B1 (de)
JP (1) JP5214375B2 (de)
CN (1) CN101382297B (de)
CA (1) CA2639356C (de)
DE (1) DE602008001042D1 (de)
FR (1) FR2920523B1 (de)
RU (1) RU2484377C2 (de)

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US8020385B2 (en) * 2008-07-28 2011-09-20 General Electric Company Centerbody cap for a turbomachine combustor and method
US8028529B2 (en) * 2006-05-04 2011-10-04 General Electric Company Low emissions gas turbine combustor
FR2917487B1 (fr) * 2007-06-14 2009-10-02 Snecma Sa Chambre de combustion de turbomachine a circulation helicoidale de l'air
US8584466B2 (en) * 2010-03-09 2013-11-19 Honeywell International Inc. Circumferentially varied quench jet arrangement for gas turbine combustors
EP2710298B1 (de) * 2011-05-17 2020-09-23 Safran Aircraft Engines Ringbrennkammer für ein turbinentriebwerk
WO2013023147A1 (en) * 2011-08-11 2013-02-14 Beckett Gas, Inc. Combustor
US9528698B2 (en) * 2011-08-11 2016-12-27 Richard D. Cook Burner
US10634354B2 (en) 2011-08-11 2020-04-28 Beckett Gas, Inc. Combustor
EP2808611B1 (de) * 2013-05-31 2015-12-02 Siemens Aktiengesellschaft Injektor zum Einbringen eines Brennstoff-Luft-Gemisches in eine Brennkammer
US10502425B2 (en) * 2016-06-03 2019-12-10 General Electric Company Contoured shroud swirling pre-mix fuel injector assembly
EP3517288B1 (de) 2018-01-25 2023-04-12 Korsch AG Fangschiene für eine rundlaufpresse
CN108679644A (zh) * 2018-04-02 2018-10-19 西北工业大学 一种旋流驻涡式微型燃气涡轮发动机燃烧室
US10935245B2 (en) 2018-11-20 2021-03-02 General Electric Company Annular concentric fuel nozzle assembly with annular depression and radial inlet ports
US11156360B2 (en) 2019-02-18 2021-10-26 General Electric Company Fuel nozzle 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

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Also Published As

Publication number Publication date
CA2639356C (fr) 2015-06-23
FR2920523A1 (fr) 2009-03-06
CA2639356A1 (fr) 2009-03-05
RU2484377C2 (ru) 2013-06-10
JP2009063287A (ja) 2009-03-26
JP5214375B2 (ja) 2013-06-19
US20090056338A1 (en) 2009-03-05
RU2008135874A (ru) 2010-03-10
FR2920523B1 (fr) 2009-12-18
US7614234B2 (en) 2009-11-10
CN101382297B (zh) 2011-11-23
EP2034245A1 (de) 2009-03-11
DE602008001042D1 (de) 2010-06-02
CN101382297A (zh) 2009-03-11

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