EP3267111A2 - 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
EP3267111A2
EP3267111A2 EP17175880.8A EP17175880A EP3267111A2 EP 3267111 A2 EP3267111 A2 EP 3267111A2 EP 17175880 A EP17175880 A EP 17175880A EP 3267111 A2 EP3267111 A2 EP 3267111A2
Authority
EP
European Patent Office
Prior art keywords
orifices
annular wall
rows
cooling
additional
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.)
Granted
Application number
EP17175880.8A
Other languages
English (en)
French (fr)
Other versions
EP3267111A3 (de
EP3267111B1 (de
Inventor
Matthieu François RULLAUD
Bernard Joseph Jean-Pierre Carrere
Hubert Pascal Verdier
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
Safran Helicopter Engines SAS
Original Assignee
Safran Aircraft Engines SAS
Safran Helicopter Engines SAS
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Safran Aircraft Engines SAS, Safran Helicopter Engines SAS filed Critical Safran Aircraft Engines SAS
Publication of EP3267111A2 publication Critical patent/EP3267111A2/de
Publication of EP3267111A3 publication Critical patent/EP3267111A3/de
Application granted granted Critical
Publication of EP3267111B1 publication Critical patent/EP3267111B1/de
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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 is more particularly an annular wall for direct combustion chamber or reverse flow cooled by a process known as "multiperforation".
  • annular turbomachine combustion chamber is formed of an inner annular wall (also called inner shell) and an outer annular wall (also called outer shell) which are connected upstream by a transverse wall forming chamber bottom.
  • the inner and outer shrouds are each provided with a plurality of holes and various orifices allowing air circulating around the combustion chamber to penetrate inside thereof.
  • so-called “primary” and “dilution” holes are formed in these ferrules to convey air inside the combustion chamber.
  • the air passing through the primary holes helps to create an air / fuel mixture that is burned in the chamber, while the air from the dilution holes is intended to promote the dilution of the same air / fuel mixture.
  • the inner and outer shells are subjected to the high temperatures of the gases from the combustion of the air / fuel mixture.
  • multiperforation holes are also drilled through these ferrules over their entire surface. These multiperforation orifices, generally inclined at 60 °, allow the air circulating outside the chamber to penetrate inside thereof by forming cooling air films along the shells.
  • the present invention therefore aims to overcome such drawbacks by providing an annular combustion chamber wall which provides adequate cooling of the areas directly downstream of the primary and dilution holes.
  • the presence of the additional cooling orifices arranged in an inclined manner in a plane perpendicular to the direction of flow of the combustion gases, directly downstream and as close as possible to the primary and dilution holes, makes it possible to ensure efficient cooling with respect to the classical axial multiperforation where the film of air is stopped by the presence of these holes and this without modifying the flow in the primary zone.
  • the multiperforation gyratory-axial transition zone makes it possible, by smoothing the flows, to reduce the thermal gradient at the origin of crack initiation.
  • the average temperature profile at the chamber outlet is improved because of the more efficient mixture thus obtained.
  • said inclinations are 30 ° and 60 ° respectively. 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 a row of additional orifices and a row of adjacent cooling orifices.
  • said inclinations are regularly distributed between 0 ° and 90 °.
  • 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 comprises in particular a compression section (not shown) in which air is compressed before being injected into a chamber housing 12, then into the combustion chamber 10 mounted inside thereof. Compressed air is introduced into the combustion chamber and mixed with fuel before being burned. The gases resulting from this combustion are then directed to a high-pressure turbine 14 disposed at the outlet of the combustion chamber.
  • the combustion chamber is of the annular type. It is formed of an inner annular wall 16 and an outer annular wall 18 which are joined upstream by a transverse wall 20 forming the chamber bottom. It can be direct as illustrated or reverse flow. In this case, a return bend that can also be cooled by multi-piercing is placed between the combustion chamber and the turbine distributor.
  • the inner annular walls 16 and outer 18 extend along a longitudinal axis slightly inclined relative 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, furnishes with the combustion chamber 10 annular spaces 26 into which compressed air for combustion is admitted. dilution and cooling of the chamber.
  • the inner annular walls 16 and outer 18 each have a cold side 16a, 18a disposed on the side of the annular space 26 in which the compressed air circulates and a hot side 16b, 18b turned towards the inside of the combustion chamber ( figure 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 of the previous one (the downstream means with respect to a general axial direction of flow of the gases resulting from the combustion of the air / fuel mixture inside the combustion chamber and represented by the arrow D).
  • the air that feeds the primary zone of the combustion chamber is introduced by a circumferential row of primary holes 28 formed in the inner annular walls 16 and outer 18 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 borrows a plurality of dilution holes 30 also formed in the inner annular walls 16 and outer 18 all around the circumference of these annular walls.
  • These dilution holes 30 are aligned in a circumferential row which is offset axially downstream from the rows of the primary holes 28 and they may have different diameters including alternating large and small holes. In the configuration illustrated in figure 2 these dilution holes of different diameters, however, present a downstream edge aligned on the same line 30A.
  • a plurality of cooling orifices 32 (illustrated in FIGS. Figures 2 and 3 ).
  • These orifices 32 which provide a cooling of the walls 16, 18 by multiperforation, are distributed in a plurality of rows circumferential axially spaced apart from each other. These rows of multiperforation orifices cover the entire surface of the annular walls of the chamber with the exception of specific areas which are the subject of the invention and are precisely delimited and situated between the line 28A, 30A forming an upstream transition axis and a transition axis. downstream axially offset downstream relative to this upstream axis and is substantially in front of the dilution holes (for the downstream axis 28B) is 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 be identical or not for all the rows.
  • the adjacent rows of cooling orifices are arranged so that the orifices 32 are staggered as shown in FIG. figure 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 air film thus formed flows in the direction of flow of the combustion gases inside the chamber (represented by the arrow D ).
  • the diameter d 1 of the cooling orifices 32 may be between 0.3 and 1. mm, the pitch d1 between 1 and 10 mm and their inclination ⁇ 1 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 the primary holes 28 and dilution holes 30 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 air film delivered by these additional orifices flows in a perpendicular direction due to their arrangement in a plane perpendicular to this axial direction D of flue gas flow.
  • This multiperforation carried out perpendicularly to the axis of the turbomachine (in the following description, it will speak of multiperforation gyratory as opposed to the axial multiperforation of the cooling orifices) allows to bring the additional orifices of the primary holes or dilution and therefore d 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 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 may, while remaining within the previously defined ranges of values, be substantially different from those of the cooling orifices 32, that is to say that the inclination ⁇ 2 of the additional orifices of a
  • the same row relative 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 of the cooling orifices 32.
  • the additional orifices 34 behind the row of primary holes 28 may further advantageously have different inclination, diameter, or pitch characteristics than those disposed behind the row of dilution holes. and, more particularly, within the same zone a difference of the diameter d2 and the pitch p2 can also be achieved to densify this cooling in the most thermally stressed parts, that is to say those just downstream of the holes primary and large dilution ports, when these are formed alternately of large and small orifices as illustrated in figure 2 .
  • the introduction of the gyratory multiperforation allows limiting the rise of the thermal gradient to prevent the formation of cracks downstream of the primary holes 28.
  • the multiperforation upstream of the holes of dilution 30 from the downstream transition axis 28B remaining axial type it is necessary to provide a transition zone made for example in two rows in which the additional cooling holes 34 are 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 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 so as not to lose the TuHP efficiency of the combustion chamber.
  • the average temperature profile at the chamber outlet is improved because of the more efficient mixture thus obtained.
  • This transition zone may for example be made in two rows of additional cooling holes each disposed in a plane inclined at 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 the additional holes in these inclined planes remaining unchanged.
  • this zone from the axis 30B may not exist or be integrated with the return elbow.
  • transition zone has been described at the level of the gyratory multiperforation, however, there is no prohibition to achieve it at the level of the axial multiperforation or still riding with a row of axial multiperforation inclined at 30 ° and a row of multiperforation gyratory inclined at 60 °.
  • this transition zone may comprise more than two rows, the inclination of the orifices then being evenly distributed between 0 ° (multiperforation axial) and 90 ° (multiperforation gyratory). For example, with three rows, the inclination of the orifices will be respectively 22.5 °, 45 ° and 67.5 °.
  • the flow in the primary zone is not modified, the gyration does not impact the orientation of the dilution jets and overcoming the thermal barrier allows a gain in weight and therefore cost.
  • the direction of drilling of the multiperforation gyratory is fixed by the orientation of the blades of the High Pressure distributor ( DHP) 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)
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

Related Parent Applications (2)

Application Number Title Priority Date Filing Date
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

Publications (3)

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

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

Country Status (9)

Country Link
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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US10890327B2 (en) 2018-02-14 2021-01-12 General Electric Company Liner of a gas turbine engine combustor including dilution holes with airflow features
US11255543B2 (en) 2018-08-07 2022-02-22 General Electric Company Dilution structure for gas turbine engine combustor
US11029027B2 (en) 2018-10-03 2021-06-08 Raytheon Technologies Corporation Dilution/effusion hole pattern for thick combustor panels
FR3090746B1 (fr) * 2018-12-20 2021-06-11 Safran Aircraft Engines Tuyere de post combustion comportant une chemise a perforation obliques
FR3098569B1 (fr) 2019-07-10 2021-07-16 Safran Aircraft Engines Paroi annulaire pour chambre de combustion de turbomachine comprenant des trous primaires, des trous de dilution et des orifices de refroidissement inclines
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Also Published As

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
CN203147824U (zh) 2013-08-21
EP3267111B1 (de) 2022-02-16
CN103958970B (zh) 2016-08-24
EP2771618A2 (de) 2014-09-03
WO2013060987A3 (fr) 2014-02-20
IN2014DN03138A (de) 2015-05-22

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