EP3267111B1 - Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern - Google Patents

Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern Download PDF

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Publication number
EP3267111B1
EP3267111B1 EP17175880.8A EP17175880A EP3267111B1 EP 3267111 B1 EP3267111 B1 EP 3267111B1 EP 17175880 A EP17175880 A EP 17175880A EP 3267111 B1 EP3267111 B1 EP 3267111B1
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EP
European Patent Office
Prior art keywords
orifices
annular wall
cooling
rows
air
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EP17175880.8A
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English (en)
French (fr)
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EP3267111A3 (de
EP3267111A2 (de
Inventor
Matthieu François RULLAUD
Bernard Joseph Jean-Pierre Carrere
Hubert Pascal Verdier
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Safran Aircraft Engines SAS
Safran Helicopter Engines SAS
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Safran Aircraft Engines SAS
Safran Helicopter Engines SAS
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Publication of EP3267111A2 publication Critical patent/EP3267111A2/de
Publication of EP3267111A3 publication Critical patent/EP3267111A3/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
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03042—Film cooled combustion chamber walls or domes

Definitions

  • the present invention relates to the general field of turbomachine combustion chambers. It relates more particularly to an annular wall for a direct or reverse-flow combustion chamber cooled by a so-called “multiperforation” process.
  • a turbomachine annular combustion chamber is formed of an inner annular wall (also called inner shroud) and an outer annular wall (also called outer shroud) which are connected upstream by a transverse wall forming the bottom of the chamber.
  • the inner and outer shrouds are each provided with a plurality of various holes and orifices allowing air circulating around the combustion chamber to penetrate inside the latter.
  • the inner and outer shrouds are subjected to the high temperatures of the gases resulting from the combustion of the air/fuel mixture.
  • multiperforation orifices are also drilled through these shrouds over their entire surface. These multi-perforation orifices, generally inclined at 60°, allow the air circulating outside the chamber to penetrate inside the latter by forming films of cooling air along the shrouds.
  • the object of the present invention is therefore to overcome such drawbacks by proposing an annular combustion chamber wall which ensures adequate cooling of the zones located directly downstream of the primary and dilution holes.
  • said two rows of orifices are then either two rows of additional orifices arranged immediately upstream of a row of cooling orifices, or two rows of cooling orifices arranged immediately downstream of a row of additional orifices, or else a row of additional orifices and an adjacent row of cooling orifices.
  • said geometric axes of each of said orifices are inclined respectively by 22.5°, 45° and 67.5°, with respect to a plane perpendicular to said axial direction D.
  • the direction of inclination of said additional orifices is constrained by the direction of flow of the air/fuel mixture downstream of said combustion chamber.
  • the present invention also relates to a combustion chamber and a turbomachine (having a combustion chamber) comprising an annular wall as defined above.
  • the figure 1 illustrates in its environment a combustion chamber 10 for a turbomachine.
  • a turbomachine includes in particular a compression section (not shown) in which air is compressed before being injected into a chamber casing 12, then into the combustion chamber 10 mounted inside the latter. Compressed air is introduced into the combustion chamber and mixed with fuel before being burned there. The gases resulting from this combustion are then directed towards a high-pressure turbine 14 arranged at the outlet of the combustion chamber.
  • the combustion chamber is of the annular type. It is formed of an internal annular wall 16 and an external annular wall 18 which are joined upstream by a transverse wall 20 forming the bottom of the chamber. It can be direct as shown or reverse flow where a return elbow which can also be cooled by multi-boring is placed between the combustion chamber and the turbine nozzle.
  • the inner 16 and outer 18 annular walls extend along a longitudinal axis slightly inclined with respect to the longitudinal axis 22 of the turbomachine.
  • the chamber bottom 20 is provided with a plurality of openings 20A in which fuel injectors 24 are mounted.
  • the chamber casing 12 which is formed of an inner casing 12a and an outer casing 12b, forms with the combustion chamber 10 annular spaces 26 into which is admitted compressed air intended for combustion, dilution and cooling of the chamber.
  • the internal 16 and external 18 annular walls each have a cold side 16a, 18a arranged on the side of the annular space 26 in which the compressed air circulates and a hot side 16b, 18b facing the interior of the combustion chamber ( picture 3 ).
  • the combustion chamber 10 is divided into a so-called “primary” zone (or combustion zone) and a so-called “secondary” zone (or dilution zone) located downstream from the previous one (downstream is understood in relation to a general axial flow direction of the gases resulting from the combustion of the air/fuel mixture inside the combustion chamber and materialized by the arrow D).
  • the air that supplies the primary zone of the combustion chamber is introduced through a circumferential row of primary holes 28 made in the internal 16 and external 18 annular walls of the chamber over the entire circumference of these annular walls. These primary holes have a downstream edge aligned on the same line 28A.
  • the air supplying the secondary zone of the chamber it passes through a plurality of dilution holes 30 also formed in the internal 16 and external 18 annular walls over the entire circumference of these annular walls.
  • These dilution holes 30 are aligned along a circumferential row which is offset axially downstream with respect to the rows of primary holes 28 and they can have different diameters with in particular an alternation of large and small holes. In the configuration shown in picture 2 , these dilution holes of different diameters however then have a downstream edge aligned on the same line 30A.
  • a plurality of cooling holes 32 are provided (illustrated in the figures 2 and 3 ).
  • These orifices 32 which provide cooling of the walls 16, 18 by multiperforation, are distributed according to a plurality of rows circumferential spaced axially from each other. These rows of multi-perforation orifices cover the entire surface of the annular walls of the chamber with the exception of particular zones which are the subject of the invention and which are precisely delimited and included between the line 28A, 30A forming an upstream transition axis and a transition axis downstream offset axially downstream with respect to this upstream axis and either substantially in front of the dilution holes (for the downstream axis 28B) or substantially in front of the exit plane of the chamber (for the downstream axis 30B).
  • the number and the diameter d1 of the cooling orifices 32 are identical in each of the rows.
  • the pitch p1 between two orifices of the same row is constant and may or may not be identical for all the rows.
  • the adjacent rows of cooling orifices are arranged so that the orifices 32 are staggered as shown in the picture 2 .
  • the cooling orifices 32 generally have an angle of inclination ⁇ 1 with respect to a normal N to the annular wall 16, 18 through which they are pierced.
  • This inclination ⁇ 1 allows the air passing through these orifices to form a film of air along the hot side 16b, 18b of the annular wall.
  • the inclination ⁇ 1 of the cooling orifices 32 is directed so that the film of air thus formed flows in the direction of flow of the combustion gases inside the chamber (schematized by the arrow D ).
  • the diameter d1 of the cooling orifices 32 can be between 0.3 and 1 mm, the pitch d1 comprised between 1 and 10 mm and their inclination ⁇ 1 comprised between +30° and +70°, typically +60°.
  • the primary holes 28 and the dilution holes 30 have a diameter of the order of 4 to 20 mm.
  • each annular wall 16, 18 of the combustion chamber comprises, arranged directly downstream of primary 28 and dilution 30 holes and distributed in several circumferential rows, typically at least 5 rows, from the axis of upstream transition 28A, 30A and up to the downstream transition axis 28B, 30B, a plurality of additional cooling orifices 34.
  • the film of air delivered by these additional orifices flows in a perpendicular direction due to their arrangement in a plane perpendicular to this axial direction D of flow of the combustion gases.
  • This multi-perforation made perpendicular to the axis of the turbomachine (in the rest of the description, we will speak of gyratory multi-perforation as opposed to the axial multi-perforation of the cooling orifices) makes it possible to bring the additional orifices closer to the primary or dilution holes and therefore to improve the efficiency of the air/fuel mixture.
  • the additional orifices 34 of the same row have the same diameter d2, preferably identical to the diameter d1 of the cooling orifices 32, are spaced apart by a constant pitch p2 which may or may not be identical to the pitch p1 between the cooling orifices 32 and have an inclination ⁇ 2, preferably identical to the inclination ⁇ 1 of the cooling orifices 32 but arranged in a perpendicular plane.
  • these characteristics of the additional orifices 34 can, while remaining within the ranges of values defined above, be substantially different from those of the cooling orifices 32, that is to say that the inclination ⁇ 2 of the additional orifices of a same row with respect to a normal N to the annular wall 16, 18 may be different from that ⁇ 1 of the cooling orifices, and the diameter d2 of the additional orifices of the same row may be different from that d1 of the cooling orifices 32.
  • the additional orifices 34 behind the row of primary holes 28 can also advantageously have characteristics in terms of inclination, diameter or pitch different from those arranged behind the row of dilution holes 30 and, more particularly, within the same zone a difference in the diameter d2 and the pitch p2 can also be made to densify this cooling in the most thermally stressed parts, that is to say those just downstream of the holes primary and large dilution orifices, when the latter are formed by alternating large and small orifices as illustrated in picture 2 .
  • the introduction of the gyratory multiperforation makes it possible, by limiting the rise in the thermal gradient, to avoid the formation of cracks downstream of the primary holes 28.
  • the multiperforation upstream of the dilution 30 from the downstream transition axis 28B remaining of the axial type it is necessary to provide a transition zone made in two rows in which the additional cooling holes 34 are each arranged in an inclined plane, one of 30° and the other 60° with respect to the axial direction D, the other parameters, namely the diameter d2, the pitch p2 and the inclination ⁇ 2 of these additional holes in these inclined planes remaining unchanged.
  • the introduction of the axial multiperforation makes it possible to fill the local level of gyration in order not to lose the TuHP efficiency of the combustion chamber.
  • the average temperature profile at the chamber outlet is improved due to the more effective mixing thus obtained.
  • This transition zone is made on two rows of additional cooling holes each arranged in an inclined plane, one of 30° and the other of 60° with respect to the axial direction D, the other parameters, namely the diameter d2 , the pitch p2 and the inclination ⁇ 2 of the additional holes in these inclined planes remaining unchanged.
  • this zone from axis 30B may not exist or be integrated into the return elbow.
  • transition zone has been described at the level of the gyratory multi-perforation, nothing however prohibits carrying it out at the level of the axial multi-perforation or even straddling a row of axial multi-perforation inclined at 30° and a row of gyratory multi-perforation inclined at 60°.
  • this transition zone can comprise three rows, the inclination of the orifices will then be 22.5°, 45° and 67.5° respectively.
  • the flow in the primary zone is not modified, the gyration not impacting the orientation of the dilution jets and by being freed from the thermal barrier allows a saving in mass and therefore in cost.
  • the direction of the drilling of the gyratory multiperforation is fixed by the orientation of the blades of the High Pressure distributor ( DBH) downstream of the combustion chamber.

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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)
  • Cylinder Crankcases Of Internal Combustion Engines (AREA)

Claims (6)

  1. Ringförmige Wand (16, 18) einer Brennkammer (10) einer Turbomaschine, umfassend eine kalte Seite (16a, 18a) und eine heiße Seite (16b, 18b), wobei die ringförmige Wand umfasst:
    . eine Vielzahl von Primärlöchern (28) oder Verdünnungslöchern (30), die entlang einer Umfangsreihe verteilt sind, um Luft, die auf der kalten Seite (16a, 18a) der ringförmigen Wand zirkuliert, zu ermöglichen, auf die heiße Seite (16b, 18b) zu dringen, um jeweils ein Luft/Treibstoff-Gemisch zu erzeugen bzw. die Verdünnung des Luft/Treibstoff-Gemischs sicherzustellen; und
    eine Vielzahl von Kühlöffnungen (32), um der Luft, die auf der kalten Seite (16a, 18a) der ringförmigen Wand zirkuliert, zu ermöglichen, auf die heiße Seite (16b, 18b) zu dringen, um einen Kühlluftfilm entlang der ringförmigen Wand zu bilden, wobei die Kühlöffnungen entlang einer Vielzahl von Umfangsreihen, welche axial voneinander beabstandet sind, verteilt sind, und wobei die geometrischen Achsen jeder der Kühlöffnungen in einer axialen Strömungsrichtung D der Verbrennungsgase um einen Neigungswinkel θ1 gegenüber einer Normalen N zu der ringförmigen Wand geneigt sind;
    . eine Vielzahl von zusätzlichen Kühlöffnungen (34), die direkt stromabwärts den Primärlöchern bzw. Verdünnungslöchern angeordnet und entlang einer Vielzahl von Umfangsreihen, welche axial voneinander beabstandet sind, verteilt sind,
    wobei die geometrischen Achsen jeder der zusätzlichen Kühlöffnungen in einer Ebene senkrecht zu der axialen Richtung D angeordnet und um einen Neigungswinkel θ2 gegenüber einer Normalen N zu der ringförmigen Wand geneigt sind,
    dadurch gekennzeichnet, dass sie ferner im Bereich einer Übergangszone (28B, 30B), die direkt stromabwärts der Vielzahl von Reihen von zusätzlichen Öffnungen (34) und direkt stromaufwärts der Vielzahl von Reihen von Kühlöffnungen (32) ausgebildet ist, genau zwei Reihen von Öffnungen umfasst, deren geometrische Achsen jeder der Öffnungen gegenüber einer Ebene senkrecht zu der axialen Richtung D um jeweils 30° und 60° geneigt sind.
  2. Wand nach Anspruch 1, dadurch gekennzeichnet, dass die beiden Reihen von Öffnungen zwei Reihen von zusätzlichen Öffnungen, die direkt stromaufwärts einer Reihe von Kühlöffnungen angeordnet sind, zwei Reihen von Kühlöffnungen, die direkt stromabwärts einer Reihe von zusätzlichen Öffnungen angeordnet sind, oder eine Reihe von zusätzlichen Öffnungen und eine benachbarte Reihe von Kühlöffnungen, sind.
  3. Ringförmige Wand (16, 18) einer Brennkammer (10) einer Turbomaschine, umfassend eine kalte Seite (16a, 18a) und eine heiße Seite (16b, 18b), wobei die ringförmige Wand umfasst:
    . eine Vielzahl von Primärlöchern (28) oder Verdünnungslöchern (30), die entlang einer Umfangsreihe verteilt sind, um Luft, die auf der kalten Seite (16a, 18a) der ringförmigen Wand zirkuliert, zu ermöglichen, auf die heiße Seite (16b, 18b) zu dringen, um jeweils ein Luft/Treibstoff-Gemisch zu erzeugen bzw. die Verdünnung des Luft/Treibstoff-Gemischs sicherzustellen; und
    . eine Vielzahl von Kühlöffnungen (32), um der Luft, die auf der kalten Seite (16a, 18a) der ringförmigen Wand zirkuliert, zu ermöglichen, auf die heiße Seite (16b, 18b) zu dringen, um einen Kühlluftfilm entlang der ringförmigen Wand zu bilden, wobei die Kühlöffnungen entlang einer Vielzahl von Umfangsreihen, welche axial voneinander beabstandet sind, verteilt sind, und wobei die geometrischen Achsen jeder der Kühlöffnungen in einer axialen Strömungsrichtung D der Verbrennungsgase um einen Neigungswinkel θ1 gegenüber einer Normalen N zu der ringförmigen Wand geneigt sind;
    . eine Vielzahl von zusätzlichen Kühlöffnungen (34), die direkt stromabwärts den Primärlöchern bzw. Verdünnungslöchern angeordnet und entlang einer Vielzahl von Umfangsreihen, welche axial voneinander beabstandet sind, verteilt sind,
    wobei die geometrischen Achsen jeder der zusätzlichen Kühlöffnungen in einer Ebene senkrecht zu der axialen Richtung D angeordnet und um einen Neigungswinkel θ2 gegenüber einer Normalen N zu der ringförmigen Wand geneigt sind,
    dadurch gekennzeichnet, dass sie ferner im Bereich einer Übergangszone (28B, 30B), die direkt stromabwärts der Vielzahl von Reihen von zusätzlichen Öffnungen (34) und direkt stromaufwärts der Vielzahl von Reihen von Kühlöffnungen (32) ausgebildet ist, genau drei Reihen von Öffnungen umfasst, deren geometrische Achsen jeder der Öffnungen gegenüber einer Ebene senkrecht zu der axialen Richtung D um jeweils 22,5°, 45° und 67,5° geneigt sind.
  4. Wand nach einem der Ansprüche 1 bis 3, dadurch gekennzeichnet, dass die Neigungsrichtung von zusätzlichen Öffnungen durch die Strömungsrichtung des Luft/Treibstoff-Gemisches stromabwärts der Brennkammer eingeschränkt wird.
  5. Brennkammer (10) einer Turbomaschine, die wenigstens eine ringförmige Wand (16, 18) nach einem der Ansprüche 1 bis 4 umfasst.
  6. Turbomaschine, umfassend eine Brennkammer (10), die wenigstens eine ringförmige Wand (16, 18) nach einem der Ansprüche 1 bis 4 umfasst.
EP17175880.8A 2011-10-26 2012-10-25 Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern Active EP3267111B1 (de)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR1159704A FR2982008B1 (fr) 2011-10-26 2011-10-26 Paroi annulaire de chambre de combustion a refroidissement ameliore au niveau des trous primaires et de dilution
PCT/FR2012/052446 WO2013060987A2 (fr) 2011-10-26 2012-10-25 Paroi annulaire de chambre de combustion à refroidissement amélioré au niveau des trous primaires et/ou de dilution
EP12790620.4A EP2771618B8 (de) 2011-10-26 2012-10-25 Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern

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EP12790620.4A Division EP2771618B8 (de) 2011-10-26 2012-10-25 Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern
EP12790620.4A Division-Into EP2771618B8 (de) 2011-10-26 2012-10-25 Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern

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EP3267111A2 EP3267111A2 (de) 2018-01-10
EP3267111A3 EP3267111A3 (de) 2018-02-28
EP3267111B1 true EP3267111B1 (de) 2022-02-16

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EP17175880.8A Active EP3267111B1 (de) 2011-10-26 2012-10-25 Ringförmige brennkammerwand mit verbesserter kühlung an den primär- und/oder verdünnungsluftlöchern

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Country Status (9)

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US (1) US10551064B2 (de)
EP (2) EP2771618B8 (de)
JP (1) JP6177785B2 (de)
CN (2) CN203147824U (de)
BR (1) BR112014010215A8 (de)
CA (1) CA2852393C (de)
FR (1) FR2982008B1 (de)
IN (1) IN2014DN03138A (de)
WO (1) WO2013060987A2 (de)

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Publication number Publication date
RU2014121037A (ru) 2015-12-10
WO2013060987A2 (fr) 2013-05-02
FR2982008B1 (fr) 2013-12-13
BR112014010215A2 (pt) 2017-06-13
JP6177785B2 (ja) 2017-08-09
US10551064B2 (en) 2020-02-04
US20140260257A1 (en) 2014-09-18
CN103958970A (zh) 2014-07-30
CA2852393C (fr) 2020-08-04
EP2771618B1 (de) 2017-06-14
JP2014531015A (ja) 2014-11-20
BR112014010215A8 (pt) 2017-06-20
EP3267111A3 (de) 2018-02-28
CA2852393A1 (fr) 2013-05-02
FR2982008A1 (fr) 2013-05-03
EP2771618B8 (de) 2017-08-16
EP3267111A2 (de) 2018-01-10
CN203147824U (zh) 2013-08-21
CN103958970B (zh) 2016-08-24
EP2771618A2 (de) 2014-09-03
WO2013060987A3 (fr) 2014-02-20
IN2014DN03138A (de) 2015-05-22

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