EP1818617B1 - Transverse wall of a combustion chamber equipped with multi-perforation holes - Google Patents
Transverse wall of a combustion chamber equipped with multi-perforation holes Download PDFInfo
- 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
- Authority
- EP
- European Patent Office
- Prior art keywords
- deflectors
- wall
- combustion chamber
- deflector
- turbine engine
- 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.)
- Active
Links
- 238000002485 combustion reaction Methods 0.000 title claims description 32
- 238000002347 injection Methods 0.000 claims description 25
- 239000007924 injection Substances 0.000 claims description 25
- 238000001816 cooling Methods 0.000 claims description 20
- 239000000446 fuel Substances 0.000 claims description 18
- 238000009434 installation Methods 0.000 claims 1
- 238000005259 measurement Methods 0.000 description 5
- 239000000203 mixture Substances 0.000 description 3
- 238000011144 upstream manufacturing Methods 0.000 description 3
- 239000000567 combustion gas Substances 0.000 description 2
- 230000006866 deterioration Effects 0.000 description 2
- 239000007789 gas Substances 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 238000005452 bending Methods 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 230000001627 detrimental effect Effects 0.000 description 1
- 239000007769 metal material Substances 0.000 description 1
Images
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
- F23R3/02—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.
Landscapes
- 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)
Description
La présente invention se rapporte au domaine général des chambres de combustion de turbomachine. Elle vise plus particulièrement la paroi d'une chambre de combustion annulaire qui est destinée à relier transversalement les parois longitudinales de cette même chambre.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.
Typiquement, une chambre de combustion annulaire de turbomachine est formée de deux parois annulaires longitudinales (une paroi interne et une paroi externe) qui sont reliées en amont par une paroi transversale également annulaire formant fond de chambre.Typically, an 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.
Le fond de chambre est muni d'une pluralité d'ouvertures de forme sensiblement circulaire qui sont régulièrement répartis sur toute la circonférence. Dans ces ouvertures sont montés des systèmes d'injection qui mélangent l'air et le carburant. Ce pré-mélange est destiné à être brûlé à l'intérieur de la chambre de combustion.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.
Afin de protéger le fond de chambre contre les températures très élevées des gaz issus de la combustion du mélange air/carburant dans la chambre de combustion, des déflecteurs formant écrans thermiques sont également montés dans chaque ouverture du fond de chambre autour des systèmes d'injection.In order to protect the bottom of the chamber against the very high temperatures of the gases resulting from the combustion of the air / fuel mixture in the combustion chamber, deflectors forming heat shields are also mounted in each opening of the chamber bottom around the injection systems. .
Le fond de chambre présente généralement une pluralité de trous de multiperforation qui sont percés dans les zones en regard des déflecteurs. Ces trous de multiperforation sont des passages pour de l'air destiné au refroidissement par impact des déflecteurs.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.
Par ailleurs, le fond de chambre se présente sous la forme d'un anneau sensiblement plan qui est centré sur l'axe longitudinal de la turbomachine. Celui-ci peut être, soit perpendiculaire à l'axe longitudinal de la turbomachine, soit incliné (vers l'intérieur ou l'extérieur) par rapport à cet axe.Furthermore, 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.
De même, les déflecteurs se présentent généralement sous la forme d'une plaque métallique de forme sensiblement rectangulaire qui est centrée sur l'axe de symétrie du système d'injection et qui est brasée sur le fond de chambre. Le document
Dans le cas où le fond de chambre est incliné par rapport à l'axe longitudinal de la turbomachine, il présente une forme tronconique avec l'axe de symétrie des systèmes d'injection dirigés vers l'intérieur ou l'extérieur. En fonctionnement, il en résulte que la distance séparant le fond de chambre de chaque déflecteur monté dans les ouvertures n'est pas constante lorsque l'on s'écarte de l'axe de symétrie des systèmes d'injection. Aussi, le refroidissement par multiperforation des déflecteurs n'est pas homogène, ce qui conduit à une forte détérioration des déflecteurs particulièrement préjudiciable à la durée de vie de la chambre de combustion.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.
La présente invention a donc pour but principal de pallier de tels inconvénients en proposant une paroi transversale de chambre de combustion de forme tronconique permettant d'obtenir un refroidissement efficace et homogène des déflecteurs.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.
Ce but est atteint grâce à une paroi annulaire destinée à relier transversalement des parois longitudinales d'une chambre de combustion annulaire de turbomachine, ladite paroi étant sensiblement plane, inclinée par rapport à un axe longitudinal de la turbomachine, et comportant une pluralité de déflecteurs formés chacun par une plaque plane sensiblement rectangulaire, lesdits déflecteurs étant montés sur la paroi annulaire et comportant chacun une ouverture pour le montage d'un système d'injection de carburant et une pluralité de trous de multiperforation formés en regard des déflecteurs autour de leur ouverture pour permettre un passage d'air destiné au refroidissement desdits déflecteurs, et dans laquelle, conformément à l'invention, chaque déflecteur comporte des moyens pour forcer le flux d'air de refroidissement des déflecteurs à s'écouler radialement par rapport à l'axe longitudinal de la turbomachine autour des systèmes d'injection de carburant.This object is achieved by means of an 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 by a substantially rectangular flat plate, said 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.
En créant des moyens pour forcer le flux d'air de refroidissement des déflecteurs à s'écouler radialement autour des systèmes d'injection de carburant, il est possible d'obtenir un refroidissement homogène sur toute la surface des déflecteurs. Ainsi, tout risque de détérioration des déflecteurs est évité. La durée de vie du fond de chambre s'en trouve donc augmentée.By creating means to force the flow of cooling air of the baffles to flow radially around the fuel injection systems, it is possible to obtain uniform cooling over the entire surface of the baffles. Thus, any risk of deterioration of the deflectors is avoided. The life of the chamber bottom is thus increased.
Selon un mode de réalisation de l'invention, chaque déflecteur comporte au moins deux déformations formant des chicanes pour l'écoulement du flux d'air de refroidissement, lesdites déformations s'étendant radialement par rapport à l'axe longitudinal de la turbomachine de part et d'autre de l'ouverture du déflecteur.According to one embodiment of the invention, 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.
La présence de telles chicanes permettent de guider radialement le flux d'air de refroidissement des déflecteurs autour des systèmes d'injection de carburant.The presence of such baffles makes it possible to radially guide the flow of cooling air of the baffles around the fuel injection systems.
Les déformations du déflecteur peuvent se présenter sous la forme de gorges, chaque gorge ayant une profondeur de préférence comprise entre 1 et 2 mm.Deflections of the deflector may be in the form of grooves, each groove having a depth preferably between 1 and 2 mm.
Selon un autre mode de réalisation de l'invention, la distance entre les extrémités radiales externes respectives de la paroi et des déflecteurs au niveau d'un plan radial de symétrie des déflecteurs est inférieure ou supérieure à celle au niveau des extrémités latérales des déflecteurs.According to another embodiment of the invention, 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.
La présence de ces écarts de distances au niveau des extrémités radiales externes respectives de la paroi et des déflecteurs permet également de guider le flux d'air de refroidissement autour des systèmes d'injection de carburant.The presence of these distance differences at the respective outer radial ends of the wall and the baffles also guides the flow of cooling air around the fuel injection systems.
La présente invention a également pour objet une chambre de combustion et une turbomachine munie d'une chambre de combustion comportant une paroi transversale telle que définie précédemment.The present invention also relates to a combustion chamber and a turbomachine provided with a combustion chamber having a transverse wall as defined above.
D'autres caractéristiques et avantages de la présente invention ressortiront de la description faite ci-dessous, en référence aux dessins annexés qui en illustrent un exemple de réalisation dépourvu de tout caractère limitatif. Sur les figures :
- la
figure 1 est une vue en coupe longitudinale d'une chambre de combustion de turbomachine dans son environnement ; - la
figure 2 est une vue partielle de la paroi transversale selon un mode de réalisation de l'invention ; - la
figure 3 représente des courbes montrant l'évolution de l'entrefer entre les déflecteurs et une paroi transversale ; - la
figure 4 est une vue en coupe selon IV-IV de lafigure 3 ; et - les
figures 5 et 6 sont des vues partielles de parois transversales selon un autre mode de réalisation de l'invention.
- the
figure 1 is a longitudinal sectional view of a turbomachine combustion chamber in its environment; - the
figure 2 is a partial view of the transverse wall according to one embodiment of the invention; - the
figure 3 represents curves showing the evolution of the air gap between the baffles and a transverse wall; - the
figure 4 is a sectional view along IV-IV of thefigure 3 ; and - the
Figures 5 and 6 are partial views of transverse walls according to another embodiment of the invention.
La
L'air comprimé est introduit dans la chambre de combustion et mélangé à du carburant avant d'y être brûlé. Les gaz issus de cette combustion sont alors dirigés vers une turbine haute-pression 5 disposée en sortie de la chambre de combustion 4.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-
La chambre de combustion 4 est de type annulaire. Elle est formée d'une paroi annulaire interne 6 et d'une paroi annulaire externe 8 qui sont réunies en amont (par rapport au sens d'écoulement des gaz de combustion dans la chambre de combustion) par une paroi transversale 10 formant fond de chambre.The combustion chamber 4 is of annular type. It is formed of an inner
Les parois interne 6 et externe 8 de la chambre de combustion s'étendent selon un axe longitudinal légèrement incliné par rapport à l'axe longitudinal X-X de la turbomachine. Elles peuvent être réalisés en un matériau métallique ou composite.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.
La paroi transversale 10 de la chambre de combustion est généralement obtenue par emboutissage d'une tôle métallique. Son épaisseur est typiquement de l'ordre de 1,5 mm environ.The
La paroi transversale 10 se présente sous la forme d'un anneau centré sur l'axe longitudinal X-X de la turbomachine. Elle se compose d'une partie principale 10a sensiblement plane (
Par ailleurs, la partie principale 10a de la paroi transversale est inclinée vers l'extérieur de l'anneau par rapport à l'axe longitudinal X-X de la turbomachine, c'est-à-dire que la paroi transversale a une forme sensiblement tronconique.Furthermore, the
L'invention s'applique également aux parois transversales dont la partie principale est inclinée vers l'intérieur de l'anneau (c'est-à-dire vers l'axe longitudinal X-X de la turbomachine).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).
La partie principale 10a de la paroi transversale 10 est pourvue d'une pluralité d'ouvertures 12, par exemple au nombre de dix-huit et de forme circulaire, qui sont régulièrement espacées sur toute la circonférence de la paroi transversale 10.The
Ces ouvertures 12 sont chacune destinées à recevoir un système d'injection 14 d'un mélange air/carburant. Ce dernier se compose notamment d'une buse d'injection de carburant 14a et d'un bol 14b équipé de vrilles à air.These
La buse et le bol sont centrés sur un axe de symétrie Y-Y du système d'injection 14. Etant donné que la paroi transversale 10 de la chambre de combustion est de forme tronconique, cet axe de symétrie Y-Y est incliné par rapport à l'axe longitudinal Y-Y de la turbomachine.The nozzle and the bowl are centered on an axis of symmetry YY of the
Un déflecteur 16 formant écran thermique est également monté dans chaque ouverture 12 de la paroi transversale 10 autour des systèmes d'injection 14.A
Comme représenté sur la
Une pluralité de trous de multiperforation 18 formant un maillage sont percés au travers de la paroi transversale 10 de la chambre de combustion autour de chaque ouverture 12 en regard des déflecteurs 16. Ils permettent à de l'air circulant autour de la chambre de combustion de venir refroidir par impact les déflecteurs.A plurality of multiperforation holes 18 forming a mesh are pierced through the
En fonctionnement, du fait que la paroi transversale 10 de la chambre de combustion est tronconique, il a été constaté que la distance (ou entrefer) d séparant les déflecteurs 16 de la paroi transversale n'est constante (de l'ordre de 1,5 à 4 mm) que dans le plan P passant par l'axe de symétrie Y-Y du système d'injection et l'axe longitudinal X-X de la turbomachine (aussi appelé plan radial de symétrie des déflecteurs - voir
La
Sur cette figure, la variation relative de l'entrefer est définie comme le rapport entre la mesure de l'entrefer d effectuée localement et la mesure réalisée au niveau du plan de symétrie P des déflecteurs.In this figure, 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.
De même, la position angulaire θ est définie par rapport au plan de symétrie P des déflecteurs (l'angle de 0° correspond à une mesure sur le plan de symétrie P et l'angle de 10° correspond à une mesure sur l'une des extrémités angulaires du déflecteur).Similarly, the angular position θ 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).
Les courbes R0, Rint et Rext de cette
On constate que l'entrefer d séparant la paroi transversale des déflecteurs varie fortement vers les extrémités latérales des déflecteurs. Il en résulte un mauvais refroidissement des déflecteurs.It is noted that the air gap d separating the transverse wall of the baffles varies strongly towards the lateral ends of the baffles. This results in poor cooling of the baffles.
Conformément à l'invention, des moyens sont prévus pour forcer le flux d'air de refroidissement des déflecteurs 16 à s'écouler radialement autour des systèmes d'injection de carburant 14.According to the invention, means are provided for forcing the flow of cooling air of the
Forcer le flux d'air de refroidissement des déflecteurs 16 à s'écouler radialement autour des systèmes d'injection de carburant 14 permet d'obtenir un refroidissement homogène sur toute la surface des déflecteurs.Forcing the flow of cooling air of the
Selon un premier mode de réalisation de l'invention représenté par les
Ces déformations 20 s'étendant radialement de part et d'autre de l'ouverture 17 du déflecteur pour le passage des systèmes d'injection de carburant 14. Plus précisément, elles ont une forme d'arc de cercle, s'étendent entre les extrémités radiales interne 16b et externe 16c du déflecteur et peuvent être symétriques par rapport au plan radial P de symétrie des déflecteurs.These
Les déformations 20 sont disposées de sorte que le débit d'air central s'écoulant radialement autour des systèmes d'injection de carburant et délimité latéralement par les deux déformations soit égal à la somme des débits d'air externes s'écoulant radialement entre chaque déformation et l'extrémité latérale correspondante du déflecteur 16.The
En outre, les déformations 20 sont de préférence formées dans des zones du déflecteur qui ne sont pas en regard de trous de multiperforation.In addition, the
Comme illustré sur la
Dans ce cas, l'épaisseur e des gorges 20 (
De telles déformations peuvent aussi bien s'appliquer à des parois transversales dont les trous de multiperforation 18 forment un maillage carré (les rangées de trous sont alignés dans le sens radial et tangentiel - cas de la
Les
On note g la distance entre les extrémités radiales externes 10c, 16c respectives de la paroi transversale 10 et des déflecteurs 16 qui est mesurée au niveau du plan radial de symétrie P des déflecteurs. La distance entre les extrémités radiales externes 10c, 16c respectives de la paroi transversale 10 et des déflecteurs 16 qui est mesurée au niveau des extrémités latérales des déflecteurs est notée h.The distance between the respective outer radial ends 10c, 16c of the
Chaque déflecteur 16 étant symétrique par rapport à son plan radial P de symétrie, il en est résulte que la distance notée h est identique aux deux extrémités latérales du déflecteur.Each
Dans un mode de réalisation représenté sur la
Dans un autre mode de réalisation représenté sur la
Quelque soit le mode de réalisation, un tel écart de distance entre les extrémités radiales externes respectives de la paroi transversale et des déflecteurs permet de forcer le flux d'air de refroidissement à s'écouler radialement autour des systèmes d'injection de carburant. Le rapport entre les distances g et h est de préférence compris entre 1,5 et 2.Whatever the embodiment, such a difference in distance between the respective outer radial ends of the transverse wall and the baffles makes it possible to force the flow of cooling air to flow radially around the fuel injection systems. The ratio between the distances g and h is preferably between 1.5 and 2.
On notera que la mise en oeuvre d'un tel écart de distance peut aussi bien s'appliquer aux extrémités radiales internes respectives de la paroi transversale et des déflecteurs. Ainsi, la distance entre les extrémités radiales internes respectives de la paroi et des déflecteurs au niveau du plan radial de symétrie des déflecteurs peut être inférieure ou supérieure à celle au niveau des extrémités latérales des déflecteurs.It will be noted that 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. Thus, 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.
Claims (8)
- Annular wall (10) intended to connect transversely longitudinal walls (6, 8) of an annular combustion chamber (4) of a turbine engine, said wall (10) being essentially flat, inclined in relation to a longitudinal axis (X-X) of the turbine engine, and comprising:a plurality of deflectors (16), each formed by an essentially rectangular flat sheet, said deflectors being mounted on the annular wall (10) and each comprising an aperture (17) for the installation of a fuel injection system (14); anda plurality of multi-perforation holes (18) formed in relation to the deflectors (16) around their aperture (17) so as to allow a passage of air intended for the cooling of the said deflectors;characterised in that each deflector (16) comprises means to force the flow of air for cooling the deflectors to flow radially in relation to the longitudinal axis (X-X) of the turbine engine around the fuel injection systems.
- Wall according to Claim 1, in which each deflector (16) comprises at least two deformations (20) forming chicanes for the movement of the flow of cooling air, said deformations (20) extending radially in relation to the longitudinal axis (X-X) of the turbine engine on both sides of the aperture (17) of the said deflector.
- Wall according to Claim 2, in which the deformations of the deflector are in the form of throats (20).
- Wall according to Claim 3, in which the throats (20) each have a thickness (e) of between 1 and 2mm.
- Wall according to Claim 1, in which the distance (g) between the respective external radial ends (10c, 16c) of the wall (10) and the deflectors (16) at the level of a radial plane of symmetry (P) of the deflectors is less than that (h) at the level of the lateral ends of the said deflectors.
- Wall according to Claim 1, in which the distance (g) between the respective external radial ends (10c, 16c) of the wall (10) and the deflectors (16) at the level of a radial plane of symmetry (P) of the deflectors is greater than that (h) at the level of the lateral ends of the said deflectors.
- Combustion chamber (4) of a turbine engine, comprising at least one annular wall (10) according to any one of Claims 1 to 6.
- Turbine engine comprising a combustion chamber (4) having at least one annular wall (10) according to any one of Claims 1 to 6.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR0650459A FR2897107B1 (en) | 2006-02-09 | 2006-02-09 | CROSS-SECTIONAL COMBUSTION CHAMBER WALL HAVING MULTIPERFORATION HOLES |
Publications (2)
Publication Number | Publication Date |
---|---|
EP1818617A1 EP1818617A1 (en) | 2007-08-15 |
EP1818617B1 true EP1818617B1 (en) | 2012-08-29 |
Family
ID=37101624
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP07101655A Active EP1818617B1 (en) | 2006-02-09 | 2007-02-02 | Transverse wall of a combustion chamber equipped with multi-perforation holes |
Country Status (7)
Country | Link |
---|---|
US (1) | US7992391B2 (en) |
EP (1) | EP1818617B1 (en) |
JP (1) | JP2007211774A (en) |
CN (1) | CN101016999A (en) |
CA (1) | CA2577595C (en) |
FR (1) | FR2897107B1 (en) |
RU (1) | RU2426948C2 (en) |
Families Citing this family (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR2910115B1 (en) * | 2006-12-19 | 2012-11-16 | Snecma | DEFLECTOR FOR BOTTOM OF COMBUSTION CHAMBER, COMBUSTION CHAMBER WHERE IT IS EQUIPPED AND TURBOREACTOR COMPRISING THEM |
FR2920525B1 (en) * | 2007-08-31 | 2014-06-13 | Snecma | SEPARATOR FOR SUPPLYING THE COOLING AIR OF A TURBINE |
FR2932251B1 (en) * | 2008-06-10 | 2011-09-16 | Snecma | COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE COMPRISING CMC DEFLECTORS |
US7712314B1 (en) | 2009-01-21 | 2010-05-11 | Gas Turbine Efficiency Sweden Ab | Venturi cooling system |
FR2958013B1 (en) * | 2010-03-26 | 2014-06-20 | Snecma | TURBOMACHINE COMBUSTION CHAMBER WITH CENTRIFUGAL COMPRESSOR WITHOUT DEFLECTOR |
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)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR2410138A2 (en) * | 1977-11-29 | 1979-06-22 | Snecma | COMBUSTION CHAMBERS FOR GAS TURBINE ENGINES |
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 (en) * | 1994-08-01 | 1996-02-08 | Bmw Rolls Royce Gmbh | Heat shield for a gas turbine combustor |
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 (en) * | 2002-03-07 | 2004-11-19 | Snecma Moteurs | MULTI-MODEL INJECTION SYSTEM FOR AN AIR / FUEL MIXTURE IN A COMBUSTION CHAMBER |
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 |
-
2006
- 2006-02-09 FR FR0650459A patent/FR2897107B1/en not_active Expired - Fee Related
-
2007
- 2007-02-02 EP EP07101655A patent/EP1818617B1/en active Active
- 2007-02-02 US US11/670,534 patent/US7992391B2/en active Active
- 2007-02-06 JP JP2007026385A patent/JP2007211774A/en active Pending
- 2007-02-08 CA CA2577595A patent/CA2577595C/en active Active
- 2007-02-08 RU RU2007104918/06A patent/RU2426948C2/en active
- 2007-02-09 CN CNA2007100050119A patent/CN101016999A/en active Pending
Also Published As
Publication number | Publication date |
---|---|
FR2897107A1 (en) | 2007-08-10 |
US7992391B2 (en) | 2011-08-09 |
US20070180834A1 (en) | 2007-08-09 |
EP1818617A1 (en) | 2007-08-15 |
CN101016999A (en) | 2007-08-15 |
JP2007211774A (en) | 2007-08-23 |
CA2577595C (en) | 2014-12-23 |
FR2897107B1 (en) | 2013-01-18 |
RU2426948C2 (en) | 2011-08-20 |
RU2007104918A (en) | 2008-08-20 |
CA2577595A1 (en) | 2007-08-09 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
EP1818617B1 (en) | Transverse wall of a combustion chamber equipped with multi-perforation holes | |
CA2577527C (en) | Turbine engine combustion chamber with tangential slots | |
EP1818615B1 (en) | Annular combustion chamber of a turbomachine | |
EP2815183B1 (en) | Air and fuel injection device for a turbomachine combustion chamber | |
EP1607582B1 (en) | Mounting of a gas turbine combustor with integrated turbine inlet guide conduit | |
EP2524169B1 (en) | Multi-bored combustion chamber having counter-rotating tangential flows | |
EP2678610B1 (en) | Annular combustion chamber for a turbine engine including improved dilution openings | |
EP1265035B1 (en) | Double mounting of a ceramic matrix composite combustion chamber | |
FR2897418A1 (en) | ANNULAR COMBUSTION CHAMBER OF A TURBOMACHINE | |
FR2970736A1 (en) | TRANSITION PIECE IMPACT SLEEVE FOR GAS TURBINE | |
FR2937680A1 (en) | ROTOR DEHUILEUR FOR TURBOMACHINE | |
EP3569929B1 (en) | Assembly for a turbine engine combustion chamber | |
FR2999276A1 (en) | GAS BURNER COMPRISING A BURNER HEAD | |
EP1930659B1 (en) | Jet engine combustion chamber | |
EP3683429B1 (en) | Internal structure of a primary ejection conduit | |
EP1939528B1 (en) | Deflector for the bottom of a combustion chamber, combustion chamber equipped with same and jet engine comprising them | |
FR2921462A1 (en) | ANNULAR COMBUSTION CHAMBER FOR A GAS TURBINE ENGINE | |
FR2893389A1 (en) | Partition for turbomachine combustion chamber has apertures for mounting fuel injectors and deflector plates which have microperforations for cooling air and larger perforations to increase air flow in zones at corners of plate | |
FR3087847A1 (en) | LOBE MIXER FOR PROMOTING MIXTURE OF CONFLUENT FLOWS | |
EP3928034B1 (en) | Combustion chamber for a turbomachine | |
WO2020224890A1 (en) | Turbine engine separate flow mixer | |
WO2019063935A1 (en) | Injection wall for a rocket engine combustion chamber | |
FR2896304A1 (en) | Annular wall for turbine engine combustion chamber has series of peripheral radial or tangential grooves to shift resonance frequency above combustion frequency | |
EP4042070A1 (en) | Pre-vaporisation tube for a turbine engine combustion chamber | |
FR3015384A1 (en) | ANNULAR COMBUSTION CHAMBER IN A TURBOMACHINE |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PUAI | Public reference made under article 153(3) epc to a published international application that has entered the european phase |
Free format text: ORIGINAL CODE: 0009012 |
|
17P | Request for examination filed |
Effective date: 20070202 |
|
AK | Designated contracting states |
Kind code of ref document: A1 Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU LV MC NL PL PT RO SE SI SK TR |
|
AX | Request for extension of the european patent |
Extension state: AL BA HR MK YU |
|
AKX | Designation fees paid |
Designated state(s): DE FR GB IT |
|
GRAP | Despatch of communication of intention to grant a patent |
Free format text: ORIGINAL CODE: EPIDOSNIGR1 |
|
GRAS | Grant fee paid |
Free format text: ORIGINAL CODE: EPIDOSNIGR3 |
|
GRAA | (expected) grant |
Free format text: ORIGINAL CODE: 0009210 |
|
AK | Designated contracting states |
Kind code of ref document: B1 Designated state(s): DE FR GB IT |
|
REG | Reference to a national code |
Ref country code: GB Ref legal event code: FG4D Free format text: NOT ENGLISH |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R096 Ref document number: 602007025069 Country of ref document: DE Effective date: 20121025 |
|
PLBE | No opposition filed within time limit |
Free format text: ORIGINAL CODE: 0009261 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT |
|
26N | No opposition filed |
Effective date: 20130530 |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R097 Ref document number: 602007025069 Country of ref document: DE Effective date: 20130530 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: PLFP Year of fee payment: 10 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: PLFP Year of fee payment: 11 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: PLFP Year of fee payment: 12 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: CD Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170707 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: DE Payment date: 20240123 Year of fee payment: 18 Ref country code: GB Payment date: 20240123 Year of fee payment: 18 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: IT Payment date: 20240123 Year of fee payment: 18 Ref country code: FR Payment date: 20240123 Year of fee payment: 18 |