EP1489359B1 - Ringförmige Brennkammer für eine Turbomaschine - Google Patents

Ringförmige Brennkammer für eine Turbomaschine Download PDF

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Publication number
EP1489359B1
EP1489359B1 EP04102723A EP04102723A EP1489359B1 EP 1489359 B1 EP1489359 B1 EP 1489359B1 EP 04102723 A EP04102723 A EP 04102723A EP 04102723 A EP04102723 A EP 04102723A EP 1489359 B1 EP1489359 B1 EP 1489359B1
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EP
European Patent Office
Prior art keywords
axial wall
zone
combustion chamber
perforations
section
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Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
EP04102723A
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English (en)
French (fr)
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EP1489359A1 (de
Inventor
Frédéric Bruno Beule
Michel André Albert Desaulty
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Safran Aircraft Engines SAS
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SNECMA SAS
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Publication of EP1489359A1 publication Critical patent/EP1489359A1/de
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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
    • F23R3/02—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
    • F23R3/42—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 generally to the field of turbomachine annular combustion chambers, and more particularly to that of means for thermally protecting these combustion chambers.
  • annular turbomachine combustion chamber comprises an outer axial wall and an inner axial wall, these walls being arranged coaxially and interconnected via a chamber bottom.
  • the combustion chamber is provided with injection ports each for receiving a fuel injector to allow combustion reactions within this combustion chamber. It is furthermore noted that these injectors can also make it possible to introduce at least part of the air intended for combustion, which occurs in a primary zone of the combustion chamber situated upstream of a secondary zone. said dilution zone.
  • the chamber bottom is provided with a plurality of passages for passing cooling air inside the chamber of combustion. It is indicated that these passages can be made on deflectors equipping the bottom chamber, these deflectors, also called cups or heat shields, being provided in order to generate protection against thermal radiation.
  • These passages are usually designed to allow the initiation of a cooling air film along the hot inner surface of the outer axial wall, as well as the initiation of a cooling air film along of the hot inner surface of the inner axial wall.
  • combustion chambers of this type have proved to be relatively efficient, they nevertheless have certain major drawbacks, related to the criterion of homogeneity of the axial wall temperatures.
  • the cooling air films initiated at the bottom of the chamber are of relatively poor circumferential homogeneity, particularly when the chamber bottom is provided with deflectors.
  • the characteristics of these films are largely susceptible to change over time, mainly due to the progressive deformation of the constituent elements of the chamber bottom.
  • the tests carried out also made it possible to detect the fact that the appearance of such hot parietal zones resulted largely from the entrapment of the cooling air films, initiated from the chamber bottom, between the axial wall concerned and the layer of cooling air from the multiperforation on the same wall.
  • the invention therefore aims to provide a turbomachine annular combustion chamber, at least partially overcoming the disadvantages mentioned above relating to the embodiments of the prior art.
  • the object of the invention is to present an annular turbomachine combustion chamber, the design of which makes it possible in particular to obtain axial wall temperatures more homogeneous than those encountered in the embodiments of the prior art.
  • the subject of the invention is an annular turbomachine combustion chamber comprising an outer axial wall, an inner axial wall and a chamber bottom connecting the axial walls, the chamber bottom being provided on the one hand with a plurality of injection ports for at least fuel injection within the combustion chamber and passages for at least initiation of a cooling air film.
  • the outer and inner axial walls being multi-perforated to allow the reinforcement cooling air films.
  • each of the outer and inner axial walls is provided, in an upstream portion, with a first zone of perforations formed so that cooling air is introduced against the current inside the the combustion chamber.
  • the specific design of the combustion chamber according to the invention makes it possible to obtain very homogeneous axial wall temperatures, by allowing a particularly large fattening of the cooling air films initiated from the chamber bottom, this fattening being carried out near the latter.
  • each perforation of the first zone of the outer axial wall is made in such a way that in axial half-section the the value of the angle formed between a tangential local direction of the outer axial wall in this half-section, and a main direction of the perforation in this same half-section is between about 30 ° and 45 °.
  • each perforation of the first zone of the inner axial wall is made in such a way that, in axial half-section, the value of the angle formed between a tangential local direction of the internal axial wall in this half -section, and a main direction of the perforation in this same half-section, is between about 30 ° and 45 °.
  • each of the outer and inner axial walls is provided, downstream of the first zone of perforations, with a second zone of perforations formed so that cooling air is introduced co-current to the inside the combustion chamber.
  • each of the external and internal axial walls is provided, between the first zone and the second perforation zone, with a transient zone of perforations intended to ensure a gradual change in the direction of rotation. introduction of the cooling air inside the combustion chamber.
  • the chamber bottom has a wall between head, it can be provided that it has, upstream to downstream, a first zone of perforations made so that cooling air is introduced against the current at the interior of the combustion chamber, a transient zone of perforations, and a second zone of perforations made so that cooling air is introduced cocurrently inside this chamber of combustion.
  • the chamber is designed so that the external and internal axial walls each comprise a plurality of primary orifices and dilution orifices, a local zone of perforations formed so that cooling air is introduced locally countercurrently within the combustion chamber is then provided downstream of each of these primary orifices, and downstream of each of these dilution orifices.
  • the presence of these local areas of perforations makes it possible to eliminate the hot spots previously encountered, downstream of each of the primary and dilution orifices.
  • annular combustion chamber 1 of a turbomachine there is partially shown an annular combustion chamber 1 of a turbomachine, according to a preferred embodiment of the present invention.
  • the combustion chamber 1 comprises an external axial wall 2, as well as an internal axial wall 4, these two walls 2 and 4 being disposed coaxially along a principal longitudinal axis 6 of the chamber 1, this axis 6 also corresponding to the axis longitudinal main of the turbomachine.
  • the axial walls 2 and 4 are connected to each other via a chamber bottom 8, which in the preferred embodiment described comprises a pilot head 10 and a take-off head 12.
  • the takeoff head 12 is offset axially downstream and radially outwardly relative to the pilot head 10.
  • these heads 10 and 12, connected to each other via a head wall 19, are respectively provided with a deflector 14 and a deflector 16.
  • this chamber bottom 8 could also have other designs known to those skilled in the art, such as a design in which it does not include a deflector, without departing from the scope of the invention.
  • Each of these injection ports 18 is designed so as to cooperate with a fuel injector 20, in order to allow the combustion reactions inside this combustion chamber 1 (the injection orifices 18 of the deflectors 14 and 16 being arranged in staggered rows, only an injection port 18 and an injector 20 of the takeoff head 12 are shown in the axial half-sectional view of Figure 1).
  • these injectors 20 are also designed to allow the introduction of at least a portion of the air for combustion, the latter occurring in a primary zone 22 located in an upstream part of the chamber
  • the air intended for combustion can also be introduced inside the chamber 1 via primary orifices 24, located all around the external axial walls 2 and 4.
  • the primary orifices 24 are arranged upstream of a plurality of dilution orifices 26, the latter also being placed all around the external axial walls 2 and internal 4, and having main function of allowing the supply of air to a dilution zone 28 located downstream of the primary zone 22.
  • a cooling air flow D serving mainly to cool the hot inner surfaces 30 and 32 of the external axial walls 2 and internal 4.
  • the deflector 14 of the pilot head 10 has a passage 34 allowing the introduction of a portion of the cooling air flow D inside the combustion chamber 1, near the inner axial wall 4.
  • the passage 34 then authorizes the initiation of a cooling air film D1 along the hot inner surface 32 of the inner axial wall 4.
  • the deflector 16 of the take off head 12 has a passage 36 allowing the introduction of another part of the cooling air flow D inside the combustion chamber 1, near the wall In such a configuration, the passage 36 therefore allows the initiation of a cooling air film D2 along the hot inner surface 30 of the outer axial wall 2.
  • the outer axial walls 2 and inner 4 are each of the multiperforated type over substantially their entire length.
  • these walls 2 and 4 have a multitude of perforations 38, preferably each cylindrical of circular section, and with a diameter of between about 0.3 and 0.6 mm.
  • the perforations 38 are distributed all around the axial wall concerned, and substantially all along the same axial wall.
  • the inner axial wall 4 has a first zone 40 of perforations 38.
  • This first zone 40 consisting of circumferential rows of perforations 38 situated furthest upstream of the wall 4, is designed so that cooling air is introduced countercurrently into the cooling chamber 1 to enrich the cooling air film D1 from the chamber bottom 8.
  • each perforation 38 of the first zone 40 in axial half-section as shown in the single figure, the value of the angle A2 formed between a tangential local direction 42 of the inner axial wall 4 in this half-section. and a main direction 44 of the perforation 38 in this same half-section is between about 30 ° and 45 °.
  • each perforation 38 can be defined as forming an angle, with the inner axial wall 4, between about 30 ° and 45 °.
  • the first zone 40 consists of a number of circumferential rows of perforations 38 between one and ten, these rows corresponding to the first upstream rows of the inner axial wall 4.
  • a second zone 46 of perforations 38 Downstream of the first zone 40 of perforations 38 is a second zone 46 of perforations 38 formed so that air coolant is introduced co-currently into the combustion chamber 1.
  • each perforation 38 is made in such a way that in axial half-section, the value of the angle A4 formed between a tangential local direction 48 of the internal axial wall 4 in this half-section, and a main direction 50 of the perforation 38 in this same half-section is between about 20 ° and 90 °.
  • each perforation 38 can be defined as forming an angle, with the inner axial wall 4, between about 20 ° and 90 °.
  • the second zone 46 which is in the form of a plurality of circumferential rows of perforations 38, extends substantially to a downstream end of the inner wall 4.
  • first and second zones 42 and 46 of the inner axial wall 4 are separated by a transient zone 52 of perforations 38, these being made in such a way that their inclinations make it possible to progressively pass, upstream downstream of a flow of countercurrent cooling air to a co-current cooling air flow.
  • the transition zone 52 consists of a number of circumferential rows of perforations 38 between one and three.
  • the inclination of the perforations 38 of this zone of transition 52 could then vary progressively, from upstream to downstream, from -30 ° to 30 °.
  • the outer axial wall 2 has a first zone 54 of perforations 38.
  • This first zone 54 consisting of circumferential rows of perforations 38 situated furthest upstream of the wall 2, is designed such that cooling air is introduced countercurrently into the cooling chamber 1 to enrich the cooling air film D2 from the chamber bottom 8.
  • the value of the angle A1 formed between a tangential local direction 56 of the outer axial wall 2 in this half-section. and a main direction 58 of the perforation 38 in the same half-section is between about 30 ° and 45 °.
  • the first zone 54 consists of a number of circumferential rows of perforations 38 between one and ten, these rows also corresponding to the first upstream rows of the outer axial wall 2.
  • each perforation 38 is made in such a way that, in axial half-section, the value of the angle A3 formed between a tangential local direction 62 of the external axial wall 2 in this half-section, and a main direction 64 of the perforation 38 in this same half-section is between about 20 ° and 90 °.
  • the second zone 60 which is in the form of a plurality of circumferential rows of perforations 38, extends substantially to a downstream end of the inner wall 4.
  • first and second zones 54 and 60 of the outer axial wall 2 are also separated by a transient zone 66 of perforations 38, these being made in such a way that their inclinations make it possible to progressively pass, from upstream to downstream, from a flow of countercurrent cooling air to a co-current cooling air flow.
  • the transition zone 66 consists of a number of circumferential rows of perforations 38 between one and three.
  • the inclination of the perforations 38 of this transition zone 66 could then vary progressively, from upstream to downstream, from -30 ° to 30 °.
  • tangential local direction may correspond to a substantially parallel line to the two straight portions symbolizing the wall in the axial half-section, close to the perforation concerned.
  • main direction of the perforation can correspond to a line substantially parallel to the two line segments symbolizing the perforation concerned, always in this same axial half-section.
  • main directions of the perforations 38 correspond respectively to their main axes, in the case where these perforations 38 are traversed diametrically by the section plane.
  • a local zone 70 of perforations 38 is formed downstream of each of the primary orifices 24 and the dilution orifices 26. These local zones 70 are provided so that cooling air is introduced locally against In this way, the perforations 38 of these local zones 70 are made in substantially the same manner as that explained above for the perforations 38 of the first zones 40 and 54.
  • the local zones 70 do not extend all around the axial walls 2 and 4, but only over a circumferential length restraint.
  • the local areas 70 are not necessarily followed, downstream, by transient zones making it possible to progressively rectify the direction introducing the cooling air into the combustion chamber 1.
  • each local zone 70 of perforations 38 extends circumferentially over a length comprised between one to two times the diameter of the primary orifice 24 or the dilution orifice 26 downstream of which it is, and each of these local areas 70 has a number of rows of perforations 38 between one and five.
  • the deflector 14 of the pilot head 10 has a passage 72 allowing the introduction of a portion of the cooling air flow D inside the combustion chamber 1 , near the wall between the head 19.
  • the passage 72 then allows the initiation of a cooling air film D3 along the hot inner surface 74 of the inter-head wall 19, the latter extending mainly axially.
  • this inter-head wall 19 is also of the multi-perforated type.
  • the inter-head wall 19 has, from upstream to downstream, a first zone 76 of perforations 38 made so that air cooling device is introduced against the current inside the combustion chamber 1, a transitional zone 78 of perforations 38, and a second zone 80 of perforations 38 made in such a way that cooling is introduced co-currently into the combustion chamber 1.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Combustion Of Fluid Fuel (AREA)
  • Gas Burners (AREA)

Claims (10)

  1. Ringförmige Brennkammer (1) für eine Turbomaschine, wobei diese Kammer (1) eine sich axial erstreckende Außenwand (2), eine sich axial erstreckende Innenwand (4) und einen Kammerboden (8) aufweist, der diese sich axial erstreckenden Wände (2, 4) miteinander verbindet, wobei der Kammerboden (8) einerseits mit einer Vielzahl von Einspritzöffnungen (18) versehen ist, die dazu vorgesehen sind, zumindest das Einspritzen des Brennstoffs ins Innere der Brennkammer (1) zu ermöglichen, und andererseits mit Durchlässen (34, 36, 72) versehen ist, die die Erzeugung eines Kühlungsluftfilms (D2) entlang der heißen Innenfläche (30) der sich axial erstreckenden Außenwand (2) sowie die Erzeugung eines Kühlungsluftfilms (D1) entlang der heißen Innenfläche (32) der sich axial erstreckenden Innenwand (4) ermöglichen, wobei diese sich axial erstreckende Außenwand und Innenwand (2, 4) über im Wesentlichen ihre gesamte Länge mehrfach durchbohrt sind, so dass die Kühlungsluftfilme (D1, D2) verstärkt werden können,
    dadurch gekennzeichnet,
    dass jede dieser sich axial erstreckenden Wände, die Außenwand (2) und die Innenwand (4), in einem vorderen Teil mit einer ersten Zone (54, 40) mit Durchbohrungen (38) versehen ist, die dergestalt ausgeführt sind, dass die Kühlungsluft als Gegenströmung ins Innere der Brennkammer (1) eingeführt wird.
  2. Ringförmige Brennkammer (1) nach Anspruch 1,
    dadurch gekennzeichnet,
    dass jede Durchbohrung (38) der ersten Zone (54) der sich axial erstreckenden Außenwand (2) dergestalt ausgeführt ist, dass im axialen halben Schnitt der Wert des Winkels (A1), der zwischen einer stellenweisen tangentialen Richtung (56) der sich axial erstreckenden Außenwand (2) in diesem halben Schnitt und einer Hauptrichtung (58) der Durchbohrung (38) in diesem gleichen halben Schnitt besteht, ca. 30° bis 45° beträgt, und dadurch, dass jede Durchbohrung (38) der ersten Zone (40) der sich axial erstreckenden Innenwand (4) dergestalt ausgeführt ist, dass im axialen halben Schnitt der Wert des Winkels (A2), der zwischen einer stellenweisen tangentialen Richtung (42) der sich axial erstreckenden Innenwand (4) in diesem halben Schnitt und einer Hauptrichtung (44) der Durchbohrung (38) in diesem gleichen halben Schnitt besteht, ca. 30° bis 45° beträgt.
  3. Ringförmige Brennkammer (1) nach Anspruch 1 oder Anspruch 2,
    dadurch gekennzeichnet,
    dass die erste Zone (54, 40) mit Durchbohrungen (38) der genannten sich axial erstreckenden Außenwand (2) und Innenwand (4) jeweils eine Anzahl, und zwar von einer bis zehn außen umlaufenden Reihen aufweisen.
  4. Ringförmige Brennkammer (1) nach einem der vorhergehenden Ansprüche,
    dadurch gekennzeichnet,
    dass die sich axial erstreckende Außenwand (2) und Innenwand (4) jeweils hinter der ersten Zone (54, 40) mit Durchbohrungen (38) mit einer zweiten Zone (60, 46) mit Durchbohrungen (38) versehen ist, die dergestalt ausgeführt sind, dass die Kühlungsluft als Mitströmung ins Innere der Brennkammer (1) eingeführt wird.
  5. Ringförmige Brennkammer (1) nach Anspruch 4,
    dadurch gekennzeichnet,
    dass jede Durchbohrung (38) der zweiten Zone (60) der sich axial erstreckenden Außenwand (2) dergestalt ausgeführt ist, dass im axialen halben Schnitt der Wert des Winkels (A3), der zwischen einer stellenweisen tangentialen Richtung (62) der sich axial erstreckenden Außenwand (2) in diesem halben Schnitt und einer Hauptrichtung (64) der Durchbohrung (38) in diesem gleichen halben Schnitt besteht, ca. 20° bis 90° beträgt, und dadurch, dass jede Durchbohrung (38) der zweiten Zone (46) der sich axial erstreckenden Innenwand (4) dergestalt ausgeführt ist, dass im axialen halben Schnitt der Wert des Winkels (A4), der zwischen einer stellenweisen tangentialen Richtung (48) der sich axial erstreckenden Innenwand (4) in diesem halben Schnitt und einer Hauptrichtung (50) der Durchbohrung (38) in diesem gleichen halben Schnitt besteht, ca. 20° bis 90° beträgt.
  6. Ringförmige Brennkammer (1) nach Anspruch 5,
    dadurch gekennzeichnet,
    dass diese sich axial erstreckende Außenwand (2) und Innenwand (4) zwischen der ersten Zone (54, 40) und der zweiten Zone (60, 46) mit Durchbohrungen (38) jeweils mit einer Übergangszone (66, 52) mit Durchbohrungen (38) versehen ist.
  7. Ringförmige Brennkammer (1) nach Anspruch 6,
    dadurch gekennzeichnet,
    dass die Übergangszone (66, 52) mit Durchbohrungen (38) dieser sich axial erstreckenden Außenwand (2) und Innenwand (4) jeweils eine Anzahl von einer bis drei außen umlaufenden Reihen aufweisen.
  8. Ringförmige Brennkammer (1) nach einem der vorhergehenden Ansprüche,
    dadurch gekennzeichnet,
    dass der Kammerboden (8) eine Zwischenkopfwand (19) enthält, die von vom nach hinten einen ersten Bereich (76) mit Durchbohrungen (38), die dergestalt ausgeführt sind, dass die Kühlungsluft als Gegenströmung ins Innere der Brennkammer (1) eingeführt wird, einen Übergangsbereich (78) mit Durchbohrungen (38) sowie einen zweiten Bereich (80) mit Durchbohrungen (38), die dergestalt ausgeführt sind, dass die Kühlungsluft als Mitströmung ins Innere der Brennkammer (1) eingeführt wird, umfasst.
  9. Ringförmige Brennkammer (1) nach einem der vorhergehenden Ansprüche,
    dadurch gekennzeichnet,
    dass die sich axial erstreckende Außenwand (2) und Innenwand (4) jeweils eine Vielzahl von Primäröffnungen (24) und Verdünnungsöffnungen (26) aufweisen, wobei ein lokaler Bereich (70) mit Durchbohrungen (38), welche dergestalt ausgeführt sind, dass die Kühlungsluft lokal begrenzt als Gegenströmung ins Innere der Brennkammer (1) eingeführt wird, jeweils hinter diesen Primäröffnungen (24) sowie jeweils hinter diesen Verdünnungsöffnungen (26) vorgesehen ist.
  10. Ringförmige Brennkammer (1) nach Anspruch 9,
    dadurch gekennzeichnet,
    dass sich jeder lokale Bereich (70) mit Durchbohrungen (38) auf der Außenumfangslinie über eine Länge erstreckt, die minimal das Einfache und maximal das Zweifache des Durchmessers der Primäröffnung (24) bzw. der Verdünnungsöffnung (26), hinter der er sich befindet, beträgt.
EP04102723A 2003-06-17 2004-06-15 Ringförmige Brennkammer für eine Turbomaschine Expired - Lifetime EP1489359B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0350226A FR2856468B1 (fr) 2003-06-17 2003-06-17 Chambre de combustion annulaire de turbomachine
FR0350226 2003-06-17

Publications (2)

Publication Number Publication Date
EP1489359A1 EP1489359A1 (de) 2004-12-22
EP1489359B1 true EP1489359B1 (de) 2006-05-03

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US (1) US7155913B2 (de)
EP (1) EP1489359B1 (de)
CA (1) CA2470928C (de)
DE (1) DE602004000789T2 (de)
ES (1) ES2262094T3 (de)
FR (1) FR2856468B1 (de)
RU (1) RU2342602C2 (de)

Families Citing this family (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7506511B2 (en) * 2003-12-23 2009-03-24 Honeywell International Inc. Reduced exhaust emissions gas turbine engine combustor
US7954326B2 (en) * 2007-11-28 2011-06-07 Honeywell International Inc. Systems and methods for cooling gas turbine engine transition liners
GB2460403B (en) * 2008-05-28 2010-11-17 Rolls Royce Plc Combustor Wall with Improved Cooling
US7874157B2 (en) * 2008-06-05 2011-01-25 General Electric Company Coanda pilot nozzle for low emission combustors
US8220269B2 (en) * 2008-09-30 2012-07-17 Alstom Technology Ltd. Combustor for a gas turbine engine with effusion cooled baffle
FR2974162B1 (fr) * 2011-04-14 2018-04-13 Safran Aircraft Engines Virole de tube a flamme dans une chambre de combustion de turbomachine
FR2979416B1 (fr) * 2011-08-26 2013-09-20 Turbomeca Paroi de chambre de combustion
US10260748B2 (en) 2012-12-21 2019-04-16 United Technologies Corporation Gas turbine engine combustor with tailored temperature profile
US10634351B2 (en) * 2013-04-12 2020-04-28 United Technologies Corporation Combustor panel T-junction cooling
US10816201B2 (en) 2013-09-13 2020-10-27 Raytheon Technologies Corporation Sealed combustor liner panel for a gas turbine engine
FR3011620B1 (fr) * 2013-10-04 2018-03-09 Snecma Chambre de combustion de turbomachine pourvue d'un passage d'entree d'air ameliore en aval d'un orifice de passage de bougie
US10816206B2 (en) 2013-10-24 2020-10-27 Raytheon Technologies Corporation Gas turbine engine quench pattern for gas turbine engine combustor
RU2581267C2 (ru) * 2013-11-12 2016-04-20 Федеральное государственное бюджетное научное учреждение "Всероссийский научно-исследовательский институт электрификации сельского хозяйства" (ФГБНУ ВИЭСХ) Устройство камеры сгорания с регулируемым завихрителем для микро газотурбинного двигателя, где турбиной и компрессором является турбокомпрессор от двс
US10317080B2 (en) 2013-12-06 2019-06-11 United Technologies Corporation Co-swirl orientation of combustor effusion passages for gas turbine engine combustor
EP3099976B1 (de) 2014-01-30 2019-03-13 United Technologies Corporation Kühlfluss für führungspaneel in einer gasturbinenbrennkammer
CN109578141B (zh) * 2019-01-23 2023-10-20 中国船舶重工集团公司第七0三研究所 一种可倒车燃气轮机动力涡轮的排气涡壳
CN115183273A (zh) * 2022-07-21 2022-10-14 中国航发沈阳发动机研究所 一种加力发动机燃烧室

Family Cites Families (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2974485A (en) * 1958-06-02 1961-03-14 Gen Electric Combustor for fluid fuels
DE2018641C2 (de) * 1970-04-18 1972-05-10 Motoren Turbinen Union Umkehrbrennkammer fuer gasturbinentriebwerke
GB1492049A (en) * 1974-12-07 1977-11-16 Rolls Royce Combustion equipment for gas turbine engines
FR2402068A1 (fr) * 1977-09-02 1979-03-30 Snecma Chambre de combustion anti-pollution
US4194358A (en) * 1977-12-15 1980-03-25 General Electric Company Double annular combustor configuration
US4244178A (en) * 1978-03-20 1981-01-13 General Motors Corporation Porous laminated combustor structure
GB2021204B (en) * 1978-05-20 1982-10-13 Rolls Royce Gas turbinr combustion chamber
GB2073396B (en) * 1980-03-29 1984-02-22 Rolls Royce Gas turbine combustion chambers
GB2099978A (en) * 1981-05-11 1982-12-15 Rolls Royce Gas turbine engine combustor
US4896510A (en) * 1987-02-06 1990-01-30 General Electric Company Combustor liner cooling arrangement
US5197289A (en) * 1990-11-26 1993-03-30 General Electric Company Double dome combustor
US5197278A (en) * 1990-12-17 1993-03-30 General Electric Company Double dome combustor and method of operation
US5142871A (en) * 1991-01-22 1992-09-01 General Electric Company Combustor dome plate support having uniform thickness arcuate apex with circumferentially spaced coolant apertures
US5220795A (en) * 1991-04-16 1993-06-22 General Electric Company Method and apparatus for injecting dilution air
CA2070518C (en) * 1991-07-01 2001-10-02 Adrian Mark Ablett Combustor dome assembly
US5154060A (en) * 1991-08-12 1992-10-13 General Electric Company Stiffened double dome combustor
US5321951A (en) * 1992-03-30 1994-06-21 General Electric Company Integral combustor splash plate and sleeve
US5289687A (en) * 1992-03-30 1994-03-01 General Electric Company One-piece cowl for a double annular combustor
CA2089302C (en) * 1992-03-30 2004-07-06 Joseph Frank Savelli Double annular combustor
CA2089285C (en) * 1992-03-30 2002-06-25 Stephen Winthrop Falls Segmented centerbody for a double annular combustor
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
US5331805A (en) * 1993-04-22 1994-07-26 Alliedsignal Inc. Reduced diameter annular combustor
FR2706021B1 (fr) * 1993-06-03 1995-07-07 Snecma Chambre de combustion comprenant un ensemble séparateur de gaz.
FR2712379B1 (fr) * 1993-11-10 1995-12-29 Snecma Chambre de combustion pour turbomachine munie d'un séparateur des gaz.
US5421158A (en) * 1994-10-21 1995-06-06 General Electric Company Segmented centerbody for a double annular combustor
US5623827A (en) * 1995-01-26 1997-04-29 General Electric Company Regenerative cooled dome assembly for a gas turbine engine combustor
WO1996027766A1 (de) * 1995-03-08 1996-09-12 Bmw Rolls-Royce Gmbh Axial gestufte doppelring-brennkammer einer gasturbine
FR2733582B1 (fr) * 1995-04-26 1997-06-06 Snecma Chambre de combustion comportant une multiperforation d'inclinaison axiale et tangentielle variable
US5970716A (en) * 1997-10-02 1999-10-26 General Electric Company Apparatus for retaining centerbody between adjacent domes of multiple annular combustor employing interference and clamping fits
US5974805A (en) * 1997-10-28 1999-11-02 Rolls-Royce Plc Heat shielding for a turbine combustor
FR2777634B1 (fr) * 1998-04-16 2000-05-19 Snecma Separateur pour chambre de combustion a deux tetes
US6079199A (en) * 1998-06-03 2000-06-27 Pratt & Whitney Canada Inc. Double pass air impingement and air film cooling for gas turbine combustor walls
US6155056A (en) * 1998-06-04 2000-12-05 Pratt & Whitney Canada Corp. Cooling louver for annular gas turbine engine combustion chamber
US6474070B1 (en) * 1998-06-10 2002-11-05 General Electric Company Rich double dome combustor
US6145319A (en) * 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
US6286317B1 (en) * 1998-12-18 2001-09-11 General Electric Company Cooling nugget for a liner of a gas turbine engine combustor having trapped vortex cavity
US6279323B1 (en) * 1999-11-01 2001-08-28 General Electric Company Low emissions combustor
US6260359B1 (en) * 1999-11-01 2001-07-17 General Electric Company Offset dilution combustor liner
US6655146B2 (en) * 2001-07-31 2003-12-02 General Electric Company Hybrid film cooled combustor liner
US6513331B1 (en) * 2001-08-21 2003-02-04 General Electric Company Preferential multihole combustor liner

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CA2470928C (en) 2011-11-15
US20050042076A1 (en) 2005-02-24
RU2342602C2 (ru) 2008-12-27
ES2262094T3 (es) 2006-11-16
US7155913B2 (en) 2007-01-02
CA2470928A1 (en) 2004-12-17
FR2856468B1 (fr) 2007-11-23
DE602004000789T2 (de) 2007-05-31
EP1489359A1 (de) 2004-12-22
DE602004000789D1 (de) 2006-06-08
RU2004118309A (ru) 2006-01-10
FR2856468A1 (fr) 2004-12-24

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