EP1818617B1 - Querwand einer Brennkammer, die mit Multiperforationslöchern ausgestattet ist - Google Patents

Querwand einer Brennkammer, die mit Multiperforationslöchern ausgestattet ist Download PDF

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Publication number
EP1818617B1
EP1818617B1 EP07101655A EP07101655A EP1818617B1 EP 1818617 B1 EP1818617 B1 EP 1818617B1 EP 07101655 A EP07101655 A EP 07101655A EP 07101655 A EP07101655 A EP 07101655A EP 1818617 B1 EP1818617 B1 EP 1818617B1
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EP
European Patent Office
Prior art keywords
deflectors
wall
combustion chamber
deflector
turbine engine
Prior art date
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Active
Application number
EP07101655A
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English (en)
French (fr)
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EP1818617A1 (de
Inventor
Gérard CABOCHE
Claude Gautier
Denis Sandelis
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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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
    • 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
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

  • the present invention relates to the general field of turbomachine combustion chambers. It relates more particularly to the wall of an annular combustion chamber which is intended to connect transversely the longitudinal walls of this same chamber.
  • annular turbomachine combustion chamber is formed of two longitudinal annular walls (an inner wall and an outer wall) which are connected upstream by an equally annular transverse wall forming chamber bottom.
  • the chamber bottom is provided with a plurality of substantially circular openings which are regularly distributed over the entire circumference. In these openings are mounted injection systems that mix air and fuel. This premix is intended to be burned inside the combustion chamber.
  • deflectors forming heat shields are also mounted in each opening of the chamber bottom around the injection systems. .
  • the chamber floor generally has a plurality of multiperforation holes that are drilled in the areas facing the baffles. These multiperforation holes are passages for air for impact cooling of the baffles.
  • the chamber bottom is in the form of a substantially plane ring which is centered on the longitudinal axis of the turbomachine. This may be either perpendicular to the longitudinal axis of the turbomachine, or inclined (inwardly or outwardly) relative to this axis.
  • the deflectors are generally in the form of a metal plate of substantially rectangular shape which is centered on the axis of symmetry of the injection system and which is brazed to the chamber bottom.
  • the document US 6,164,074 has such a device where the baffle is screwed onto the bottom wall of the chamber.
  • the chamber bottom In the case where the chamber bottom is inclined relative to the longitudinal axis of the turbomachine, it has a frustoconical shape with the axis of symmetry injection systems directed inward or outward. In operation, it follows that the distance separating the chamber bottom of each deflector mounted in the openings is not constant when one deviates from the axis of symmetry of the injection systems. Also, the multiperforation cooling of the baffles is not homogeneous, which leads to a sharp deterioration of the deflectors particularly detrimental to the service life of the combustion chamber.
  • the main object of the present invention is thus to overcome such disadvantages by proposing a transverse wall of a frustoconical shaped combustion chamber making it possible to obtain efficient and homogeneous cooling of the baffles.
  • annular wall intended to transversely connect longitudinal walls of an annular turbomachine combustion chamber, said wall being substantially flat, inclined with respect to a longitudinal axis of the turbomachine, and comprising a plurality of formed deflectors.
  • each baffles being mounted on the annular wall and each having an opening for mounting a fuel injection system and a plurality of multiperforation holes formed opposite the deflectors around their opening for allow an air passage for cooling said deflectors, and wherein, according to the invention, each baffle comprises means for forcing the flow of cooling air of the baffles to flow radially relative to the longitudinal axis of the turbomachine around the fuel injection systems.
  • each deflector comprises at least two deformations forming baffles for the flow of the cooling air flow, said deformations extending radially with respect to the longitudinal axis of the turbomachine from and other of the opening of the deflector.
  • baffles make it possible to radially guide the flow of cooling air of the baffles around the fuel injection systems.
  • Deflections of the deflector may be in the form of grooves, each groove having a depth preferably between 1 and 2 mm.
  • the distance between the respective outer radial ends of the wall and the baffles at a radial plane of symmetry deflectors is less than or greater than that at the lateral ends of the baffles.
  • the present invention also relates to a combustion chamber and a turbomachine provided with a combustion chamber having a transverse wall as defined above.
  • the figure 1 illustrates a combustion chamber 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 2, then into a combustion chamber 4 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 5 disposed at the outlet of the combustion chamber 4.
  • the combustion chamber 4 is of annular type. It is formed of an inner annular wall 6 and an outer annular wall 8 which are connected upstream (with respect to the flow direction of the combustion gases in the combustion chamber) by a transverse wall forming a chamber bottom .
  • the inner 6 and outer 8 walls of the combustion chamber extend along a longitudinal axis slightly inclined relative to the longitudinal axis X-X of the turbomachine. They can be made of a metallic or composite material.
  • the transverse wall 10 of the combustion chamber is generally obtained by stamping a metal sheet. Its thickness is typically of the order of about 1.5 mm.
  • the transverse wall 10 is in the form of a ring centered on the longitudinal axis XX of the turbomachine. It consists of a substantially flat main part 10a ( figure 2 ) which extends at its two free ends by parts 10b folded upstream ( figure 1 ).
  • the main portion 10a of the transverse wall is inclined outwardly of the ring relative to the longitudinal axis XX of the turbomachine, that is to say that the transverse wall has a substantially frustoconical shape.
  • the invention also applies to transverse walls whose main part is inclined towards the inside of the ring (that is to say towards the longitudinal axis X-X of the turbomachine).
  • the main part 10a of the transverse wall 10 is provided with a plurality of openings e.g. eighteen in number and circular in shape, which are evenly spaced along the entire circumference of the transverse wall 10.
  • openings 12 are each intended to receive an injection system 14 of an air / fuel mixture.
  • the latter consists in particular of a fuel injection nozzle 14a and a bowl 14b equipped with air auger.
  • the nozzle and the bowl are centered on an axis of symmetry YY of the injection system 14. Since the transverse wall 10 of the combustion chamber is of frustoconical shape, this axis of symmetry YY is inclined relative to the axis longitudinal YY of the turbomachine.
  • a baffle 16 forming a heat shield is also mounted in each opening 12 of the transverse wall 10 around the injection systems 14.
  • the deflectors 16 are flat plates of substantially rectangular shape which each have a circular opening 17 centered on the axis of symmetry YY injection systems for the passage of the latter. They make it possible to protect the transverse wall against the high temperatures of the combustion gases.
  • a plurality of multiperforation holes 18 forming a mesh are pierced through the transverse wall 10 of the combustion chamber around each opening 12 facing the deflectors 16. They allow air circulating around the combustion chamber of come to cool by impact the deflectors.
  • the distance (or gap) d separating the baffles 16 from the transverse wall is constant (of the order of 1, 5 to 4 mm) than in the plane P passing through the axis of symmetry YY of the injection system and the longitudinal axis XX of the turbomachine (also called the radial plane of symmetry of the deflectors - see figure 2 ) and that it varies when one deviates from this radial plane of symmetry P.
  • the variation of the air gap d depends in particular on the number of injection systems equipping the combustion chamber, the height of the primary combustion zone and the mean radius of the transverse wall.
  • the figure 3 illustrates the relative variation of the air gap d as a function of the angular position ⁇ at which the measurement of the gap d is made.
  • the relative variation of the gap is defined as the ratio between the measurement of the air gap d made locally and the measurement made at the plane of symmetry P deflectors.
  • the angular position ⁇ is defined relative to the plane of symmetry P of the baffles (the angle of 0 ° corresponds to a measurement on the plane of symmetry P and the angle of 10 ° corresponds to a measurement on one angular ends of the deflector).
  • the curves R0, Rint and Rext of this figure 3 represent the relative variation of the working gap, respectively, for the mean radius 16a, for the internal radius 16b and for the outer radius 16c of the deflector 16 (these radii are shown schematically on the figure 2 ).
  • means are provided for forcing the flow of cooling air of the baffles 16 to flow radially around the fuel injection systems 14.
  • each deflector 16 has at least two deformations 20 forming baffles for the flow of the cooling air stream.
  • These deformations 20 extending radially on either side of the opening 17 of the deflector for the passage of the fuel injection systems 14. More precisely, they have a shape of an arc of a circle, extend between the internal radial ends 16b and outer 16c of the deflector and may be symmetrical with respect to the radial plane P of symmetry of the deflectors.
  • the deformations 20 are arranged so that the central air flow flowing radially around the fuel injection systems and delimited laterally by the two deformations is equal to the sum of the external air flows flowing radially between each deformation and the corresponding lateral end of the deflector 16.
  • the deformations 20 are preferably formed in areas of the baffle which are not facing multiperforation holes.
  • the deformations are advantageously in the form of grooves 20 which are for example made by stamping the deflectors 16.
  • the thickness e of the grooves 20 ( figure 2 ) can be between 1 and 2 mm.
  • the depth of the grooves is such that the distance f between the bottom of a groove 20 and the transverse wall 10 ( figure 4 ) is constant (for example of the order of 0.3 to 0.5 mm).
  • Such deformations can also be applied to transverse walls whose multiperforation holes 18 form a square mesh (the rows of holes are aligned in the radial and tangential direction - case of the figure 2 ) than transverse walls whose multiperforation holes form an equilateral grid (the holes are arranged in staggered rows with respect to each other).
  • FIGS. 5 and 6 represent another embodiment of the means for forcing the cooling air flow of the baffles to flow radially around the fuel injection systems according to the invention.
  • the distance between the respective outer radial ends 10c, 16c of the transverse wall 10 and the deflectors 16 which is measured at the level of the radial plane of symmetry P of the deflectors is noted g .
  • the distance between the respective outer radial ends 10c, 16c of the transverse wall 10 and the deflectors 16 which is measured at the lateral ends of the baffles is denoted h .
  • Each baffle 16 being symmetrical with respect to its radial plane P symmetry, it follows that the distance noted h is identical to the two lateral ends of the deflector.
  • each deflector 16 is arranged in such a way that the distance g previously defined is greater than the distance h .
  • each deflector 16 is arranged such that the distance g is smaller than the distance h . This can be obtained for example by bending the outer radial end 16c of the deflectors 16.
  • the ratio between the distances g and h is preferably between 1.5 and 2.
  • the implementation of such a difference in distance can also be applied to the respective inner radial ends of the transverse wall and the baffles.
  • the distance between the respective inner radial ends of the wall and the baffles at the radial plane of symmetry of the deflectors may be less than or greater than that at the lateral ends of the baffles.

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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 Methods Of Internal-Combustion Engines (AREA)
  • Fuel-Injection Apparatus (AREA)

Claims (8)

  1. Ringförmige Wand (10), die dazu bestimmt ist, Längswände (6, 8) einer ringförmigen Brennkammer (4) einer Turbomaschine quer zu verbinden, wobei die Wand (10) im wesentlichen eben, in bezug auf eine Längsachse (X-X) der Turbomaschine geneigt ist und umfaßt:
    eine Vielzahl von Abweisern (16), die jeweils durch eine ebene, im wesentlichen rechteckige Platte gebildet sind, wobei die Abweiser an der ringförmigen Wand (10) angebracht sind und jeweils eine Öffnung (17) für die Montage eines Treibstoffeinspritzsystems (14) umfassen, und
    eine Vielzahl von Löchern einer Mehrfachlochung (18), die gegenüber den Abweisern (16) um deren Öffnung (17) herum ausgebildet sind, um einen für die Kühlung der Abweiser bestimmten Luftdurchgang zu ermöglichen,
    dadurch gekennzeichnet, daß jeder Abweiser (16) Mittel umfaßt, um den Luftstrom zum Kühlen der Abweiser dazu zu bringen, in bezug auf die Längsachse (X-X) der Turbomaschine radial um die Treibstoffeinspritzsysteme zu strömen.
  2. Wand nach Anspruch 1, wobei jeder Abweiser (16) wenigstens zwei Verformungen (20) umfaßt, die Leitbleche für die Strömung des Kühlluftstroms bilden, wobei die Verformungen (20) in bezug auf die Längsachse (X-X) der Turbomaschine radial auf beiden Seiten der Öffnung (17) des Abweisers verlaufen.
  3. Wand nach Anspruch 2, wobei die Verformungen des Abweisers in Form von Nuten (20) vorliegen.
  4. Wand nach Anspruch 3, wobei die Nuten (20) jeweils eine Dicke (e) zwischen 1 und 2 mm aufweisen.
  5. Wand nach Anspruch 1, wobei der Abstand (g) zwischen den jeweiligen äußeren radialen Enden (10c, 16c) der Wand (10) und der Abweiser (16) im Bereich einer radialen Symmetrieebene (P) der Abweiser kleiner ist als der (h) im Bereich der Seitenenden der Abweiser.
  6. Wand nach Anspruch 1, wobei der Abstand (g) zwischen den jeweiligen äußeren radialen Enden (10c, 16c) der Wand (10) und der Abweiser (16) im Bereich einer radialen Symmetrieebene (P) der Abweiser größer ist als der (h) im Bereich der Seitenenden der Abweiser.
  7. Brennkammer (4) einer Turbomaschine, die wenigstens eine ringförmige Wand (10) nach einem der Ansprüche 1 bis 6 umfaßt.
  8. Turbomaschine, die eine Brennkammer (4) mit wenigstens einer ringförmigen Wand (10) nach einem der Ansprüche 1 bis 6 umfaßt.
EP07101655A 2006-02-09 2007-02-02 Querwand einer Brennkammer, die mit Multiperforationslöchern ausgestattet ist Active EP1818617B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0650459A FR2897107B1 (fr) 2006-02-09 2006-02-09 Paroi transversale de chambre de combustion munie de trous de multiperforation

Publications (2)

Publication Number Publication Date
EP1818617A1 EP1818617A1 (de) 2007-08-15
EP1818617B1 true EP1818617B1 (de) 2012-08-29

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EP07101655A Active EP1818617B1 (de) 2006-02-09 2007-02-02 Querwand einer Brennkammer, die mit Multiperforationslöchern ausgestattet ist

Country Status (7)

Country Link
US (1) US7992391B2 (de)
EP (1) EP1818617B1 (de)
JP (1) JP2007211774A (de)
CN (1) CN101016999A (de)
CA (1) CA2577595C (de)
FR (1) FR2897107B1 (de)
RU (1) RU2426948C2 (de)

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2910115B1 (fr) * 2006-12-19 2012-11-16 Snecma Deflecteur pour fond de chambre de combustion, chambre de combustion en etant equipee et turboreacteur les comportant
FR2920525B1 (fr) * 2007-08-31 2014-06-13 Snecma Separateur pour alimentation de l'air de refroidissement d'une turbine
FR2932251B1 (fr) * 2008-06-10 2011-09-16 Snecma Chambre de combustion de moteur a turbine a gaz comportant des deflecteurs en cmc
US7712314B1 (en) 2009-01-21 2010-05-11 Gas Turbine Efficiency Sweden Ab Venturi cooling system
FR2958013B1 (fr) * 2010-03-26 2014-06-20 Snecma Chambre de combustion de turbomachine a compresseur centrifuge sans deflecteur
US9377198B2 (en) 2012-01-31 2016-06-28 United Technologies Corporation Heat shield for a combustor
GB2543803B (en) * 2015-10-29 2019-10-30 Rolls Royce Plc A combustion chamber assembly
US11313560B2 (en) 2018-07-18 2022-04-26 General Electric Company Combustor assembly for a heat engine
US11391461B2 (en) * 2020-01-07 2022-07-19 Raytheon Technologies Corporation Combustor bulkhead with circular impingement hole pattern

Family Cites Families (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2410138A2 (fr) * 1977-11-29 1979-06-22 Snecma Perfectionnements aux chambres de combustion pour moteur a turbine a gaz
US4934145A (en) * 1988-10-12 1990-06-19 United Technologies Corporation Combustor bulkhead heat shield assembly
GB9018014D0 (en) * 1990-08-16 1990-10-03 Rolls Royce Plc Gas turbine engine combustor
GB2247522B (en) * 1990-09-01 1993-11-10 Rolls Royce Plc Gas turbine engine combustor
US5419115A (en) * 1994-04-29 1995-05-30 United Technologies Corporation Bulkhead and fuel nozzle guide assembly for an annular combustion chamber
DE4427222A1 (de) * 1994-08-01 1996-02-08 Bmw Rolls Royce Gmbh Hitzeschild für eine Gasturbinen-Brennkammer
US5581999A (en) * 1994-12-15 1996-12-10 United Technologies Corporation Bulkhead liner with raised lip
US6164074A (en) * 1997-12-12 2000-12-26 United Technologies Corporation Combustor bulkhead with improved cooling and air recirculation zone
US6155056A (en) * 1998-06-04 2000-12-05 Pratt & Whitney Canada Corp. Cooling louver for annular gas turbine engine combustion chamber
US6557349B1 (en) * 2000-04-17 2003-05-06 General Electric Company Method and apparatus for increasing heat transfer from combustors
US6530227B1 (en) * 2001-04-27 2003-03-11 General Electric Co. Methods and apparatus for cooling gas turbine engine combustors
US6497105B1 (en) * 2001-06-04 2002-12-24 Pratt & Whitney Canada Corp. Low cost combustor burner collar
US6546733B2 (en) * 2001-06-28 2003-04-15 General Electric Company Methods and systems for cooling gas turbine engine combustors
FR2836986B1 (fr) * 2002-03-07 2004-11-19 Snecma Moteurs Systeme d'injection multi-modes d'un melange air/carburant dans une chambre de combustion
US6792757B2 (en) * 2002-11-05 2004-09-21 Honeywell International Inc. Gas turbine combustor heat shield impingement cooling baffle
US8596071B2 (en) * 2006-05-05 2013-12-03 General Electric Company Method and apparatus for assembling a gas turbine engine

Also Published As

Publication number Publication date
EP1818617A1 (de) 2007-08-15
CA2577595A1 (fr) 2007-08-09
US7992391B2 (en) 2011-08-09
CN101016999A (zh) 2007-08-15
RU2007104918A (ru) 2008-08-20
FR2897107A1 (fr) 2007-08-10
CA2577595C (fr) 2014-12-23
JP2007211774A (ja) 2007-08-23
FR2897107B1 (fr) 2013-01-18
US20070180834A1 (en) 2007-08-09
RU2426948C2 (ru) 2011-08-20

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