EP0887515B1 - Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall - Google Patents

Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall Download PDF

Info

Publication number
EP0887515B1
EP0887515B1 EP98401558A EP98401558A EP0887515B1 EP 0887515 B1 EP0887515 B1 EP 0887515B1 EP 98401558 A EP98401558 A EP 98401558A EP 98401558 A EP98401558 A EP 98401558A EP 0887515 B1 EP0887515 B1 EP 0887515B1
Authority
EP
European Patent Office
Prior art keywords
blade
cavity
air
upstream
cooling
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.)
Expired - Lifetime
Application number
EP98401558A
Other languages
German (de)
French (fr)
Other versions
EP0887515A1 (en
Inventor
Yves Maurice Bailly
Xavier Gérard André Coudray
Mischael François Louis Derrien
Jean-Michel Roger Fougeres
Philippe Christian Pellier
Jean-Claude Christian Taillant
Thierry Henri Marcel Tassin
Christophe Bernard Texier
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
Original Assignee
SNECMA Moteurs SA
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 SNECMA Moteurs SA filed Critical SNECMA Moteurs SA
Publication of EP0887515A1 publication Critical patent/EP0887515A1/en
Application granted granted Critical
Publication of EP0887515B1 publication Critical patent/EP0887515B1/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/15—Two-dimensional spiral
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/25—Three-dimensional helical

Definitions

  • the invention relates to the blades of high pressure turbines turbomachinery.
  • the fixed and moving blades of high pressure turbines are subject to the high temperatures of the combustion gases of the combustion chamber. Also the blades of these blades are equipped with cooling devices supplied with air cooling taken from the high pressure compressor. This cooling air passes through circuits provided inside from dawn, then is discharged into the vein of hot gases flowing between the blades.
  • the cooling air enters the blades by the blade root, but in the fixed blades, the air of cooling can be introduced by a fixed vane base, either at the foot of dawn is at the head of dawn, the foot of dawn being the end of the blade closest to the axis of rotation of the turbine.
  • US-A-4 992 026 discloses a turbine blade comprising a hollow aerodynamic wall which extends radially between a foot blade and a blade head and which has a leading edge and an edge leakage, separated from each other and connected by a concave side wall or lower surface and a convex or upper side wall and comprising in addition to a cooling device provided inside said blade, supplied with cooling air from the blade root and intended for direct the cooling air against the interior surfaces of said side walls.
  • this blade comprises two radial partitions which connect said concave and convex side walls and which separate the interior of said blade into an upstream cavity located near the leading edge, a middle cavity located between said radial partitions and a downstream cavity located on the trailing edge side, the upstream cavity and the middle cavity are supplied with air by an inlet provided at the root of the blade, this air then being evacuated from said cavities by orifices provided at the head of the blade, while the downstream cavity is supplied with air by a separate inlet provided at the foot of the blade, this air then being evacuated by a plurality of slots formed in the trailing edge.
  • the internal wall of the upstream cavity comprises disruptors.
  • These disruptors can consist of ribs, spikes or bridges connecting the inner wall of the blade to the soul of the helical ramp.
  • the jacket of the middle cavity advantageously comprises a plurality of juxtaposed compartments which are powered successively by the same air flow.
  • the first compartment is supplied with air by the blade root, and the following compartments are supplied by the air flow coming from the previous compartment and having impacted the side walls of the dawn, by slots provided in the jacket walls under the protruding elements, the latter being made up of transverse ribs.
  • the helical ramp makes it possible to increase very significantly the internal exchange coefficient for cooling the blade in the leading edge area.
  • the cascade impact system placed in the cavity middle, allows to use the full potential of the cooling air before it is reintroduced into the vein.
  • the combination of these cooling technologies allows optimize the ventilation of the turbine blades by using at maximum the potential of the cooling air and having a thermal design leading to a mechanical service life optimal.
  • the design of the blade according to the invention makes it possible to reduce the ventilation flow and therefore increase the efficiency of the motor.
  • the drawing shows a moving blade 1 of a high turbine pressure which has a hollow aerodynamic wall 2, also referred to as a blade which extends radially between a blade root 3 and a blade head 4.
  • the aerodynamic wall 2 has four zones distinct: a rounded leading edge 5 intended to be placed opposite the flow of hot gases from the combustion chamber, an edge of tapered leak 6, distant from the leading edge and connected to the latter by a concave side wall 7, called the lower surface, and a side wall convex 8, called upper surface.
  • the side walls 7 and 8 are connected by two radial partitions 9 and 10 which separate the interior of the blade 1 into three cavities: a cavity upstream 11 located in the immediate vicinity of the leading edge 5, a cavity median 12 located between the two radial partitions 9 and 10 and a downstream cavity 13 located on the trailing edge side 6.
  • the downstream cavity 13 is the widest and occupies about two-thirds of the extent of dawn 1.
  • a third radial partition 14 further separates the downstream cavity 13 in an upstream part 15 and a downstream part 16 near the trailing edge 6.
  • a transverse partition 17 closes the lower end of the downstream cavity 13.
  • the upstream part 15 and the downstream part 16 communicate between them by an opening 18 formed at the foot of the third partition 14.
  • a plurality of slots 19 which connect the downstream part 16 of the downstream cavity 13 with the combustion gas stream which flows along the side walls 7 and 8 of the blade 1.
  • an orifice 20 is formed in the wall of the blade head 4 in line with the upstream cavity 11, and a second orifice 21, of oblong shape, is formed in the blade head 4, above the middle cavity 12.
  • two separate conduits 22 are formed. and 23 for supplying cooling air.
  • the first conduit 22 directly supplies cooling air to the ends lower of the upstream cavity 11 and the middle cavity 13, as well as this is shown in Figures 2 and 11, while the second leads 23 supplies cooling air to the upstream part 15 of the cavity downstream 13 in the vicinity of the blade head 4, this air having passed through the interior of the two side walls 6 and 7 consisting of double skins connected by at least straight bridges 24 of the upstream part 15, as shown in Figures 12 to 14.
  • the blade 1 is made at its aerodynamic wall hollow 2 into two half-blades subsequently joined by brazing, the cut between the two half-blades at the skeleton, or the dawn can be made in foundry.
  • the upstream cavity 11 located near the leading edge 5 is cooled by convection by through a helical ramp 30.
  • This ramp 30 can be obtained by foundry and be in one piece with a half-blade, or added in the upstream cavity 11 and brazed.
  • the helical ramp 30 shown in Figure 3 includes two nets 31a, 31b, however this ramp 30 may have only one single net or more than two, as needed.
  • the central body 32, or core, of the ramp 30 is not necessarily cylindrical, it can have an evolving section on the height in order to modulate the section as desired the section of cooling air passage in order to regulate the levels of exchange coefficient.
  • the cooling air circulates in a "worm" type cooling system that starts from the bottom 3 of dawn and ends at the head of dawn 5, from which the air is evacuated by orifice 20.
  • This system makes it possible to significantly increase the distance of the air flow and increase, at a fixed cooling rate, flow velocity relative to that obtained in a cavity purely radial.
  • disturbers 33 in the form inclined ribs are arranged either on the inner wall of the upstream cavity 11, ie on the helical ramp.
  • the disruptors can be made up of bridges 34 which connect the internal wall of the upstream cavity 11 to the core 32 of the helical ramp 30. These bridges 34 can be staggered.
  • the disturbers can be constituted by pins 35 arranged staggered or not on the wall internal of the upstream cavity 11.
  • the cooling device described above is implemented place in the upstream cavity 11 located in the immediate vicinity of the edge 5. This device could also be placed in other rooms. cavities.
  • the cooling air in this upstream cavity 11 circulates from centrifugal way, from the blade root 3 to the blade head 5. But the circuit can be reversed, especially in the fixed blades of turbine distributors, for example. Several helical ramps can also equip a cavity with reversal of the circuit cooling at the foot or at the head of the blade.
  • the central cavity 12 is cooled by convection using the cascade impact technology with cooling air introduced into the lower part of the cavity 12 from the conduit 22 formed in the blade root 3.
  • FIGS 2 and 8 to 11 show that a jacket 40 is introduced into the middle cavity 12.
  • This jacket 40 is produced by a mechanically welded assembly of a set of sheets beforehand drilled to make impact holes 41, and slots 42 or can be carried out directly in the foundry.
  • the shirt 40 is in the form of a chimney, of which two opposite side walls 43 and 44 bear on the walls internal of the radial partitions 9 and 10 and of which the two other walls 45 and 46, which include the impact holes 41 and the slots 42 are kept at a certain distance from the side walls 7 and 8 of dawn 1 by projecting elements 47, in the form of transverse ribs, formed on the walls 45 and 46 and regularly distributed between the blade root 3 and the blade head 4.
  • the internal cavity of the shirt 40 is divided into a certain number of compartments, referenced C1 to C7 in Figure 11, at by means of transverse partitions 48 arranged respectively, in starting from the blade root 3, under a couple of projecting elements 47 and separated from these projecting elements 47 by two facing slots 42 walls 7 and 8 of the blade 1.
  • the upper partition 48a is separated from the wall forming the blade head 4, so that the cooling air evacuated from the cavity C7 can be evacuated through the orifice 21.
  • the cooling circuit in the middle cavity 12 is carried out as follows
  • the air is brought through line 22 into compartment C1 of the jacket 40, then is evacuated from compartment C1 through the orifices impact 41, in order to strike the internal walls of the lower surface 7 and the upper surface 8 of the blade 1 in the vicinity of the blade root 3.
  • air is introduced into the second compartment C2 through the first slots 42, then discharged through the impact orifices 21 of the compartment C2 to be then reintroduced into the third compartment C3.
  • the air flows in this way to the upper compartment C7, from where it impacts the internal walls of the lower surface 7 and the upper surface 8 neighborhood of the blade head 4, then is evacuated out of the blade L by orifice 21.
  • the number of compartments can be different from 7, and the number of impact orifices 41 may be different from compartment to the other.
  • the shirt 40 described above could also be mounted in a cavity near the leading edge or the trailing edge. She can be adapted to both fixed and mobile blades. For fixed blades, the feed can be done by the blade head 4, and compartments C1 to C7 can be arranged radially, as in the example described above, or be arranged axially from the leading edge 5 to the trailing edge 6 or vice versa. This device can be applied as well for distributed impact (several rows of orifices) only for concentrated impact (a single row of holes 41).
  • the lower surface 7 and the upper surface 8 comprise at the level of the upstream part 15 of the cavity downstream 13 of the double skins 7a, 7b and 8a, 8b, connected by bridges 24.
  • the internal skins 7b, 8b are connected in the vicinity of the blade root 3 by the transverse partition 17. These two internal skins 7b, 8b extend to the vicinity of the partition forming the blade head 4, while reserving passages 50a, 50b near the blade head 4 by which, the air introduced through the orifice 23 of the blade root 3, and having circulated centrifugally between the skins 7a, 7b of the lower surface 7 and the skins 8a, 8b of the upper surface 8, is evacuated in the upstream part 15 of the downstream cavity.
  • This cooling air circulates centripetally in this upstream part 15, then enters the downstream part 16 by the opening 18. The air finally rises centrifugally in the downstream part 16 and is discharged into the stream of hot gases through the slots 19 formed in the trailing edge 6.
  • the cooling air introduced by orifice 23 is divided into two flows B1 and B2 by the partition transverse 17. These two flows B1 and B2 circulate so centrifugal through the multitude of bridges 24. These bridges 24 are obtained in foundry during casting. These bridges 24 can be staggered (see Figure 13) or arranged in a row (see figure 14).
  • the shape of the bridges can be any, of section cylindrical, square, oblong .... This device can also be used for cooling areas extending to the edge attack.
  • the constitution of the internal cooling circuits is realizes by assembling the added parts, helical ramp 30 and 40 welded shirt, in one of the half-blades, then in bringing the other half-dawn over the previous one and then brazing all the parts.
  • the cooling circuits can also be carried out entirely or partially directly in foundry.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

L'invention concerne les aubages des turbines à haute pression des turbomachines.The invention relates to the blades of high pressure turbines turbomachinery.

Les aubes fixes et mobiles des turbines à haute pression sont soumises aux températures élevées des gaz de combustion de la chambre de combustion. Aussi les pales de ces aubes sont équipées de dispositifs de refroidissement alimentés avec un air de refroidissement prélevé au niveau du compresseur à haute pression. Cet air de refroidissement transite par des circuits prévus à l'intérieur de l'aube, puis est évacué dans la veine de gaz chauds circulant entre les aubes.The fixed and moving blades of high pressure turbines are subject to the high temperatures of the combustion gases of the combustion chamber. Also the blades of these blades are equipped with cooling devices supplied with air cooling taken from the high pressure compressor. This cooling air passes through circuits provided inside from dawn, then is discharged into the vein of hot gases flowing between the blades.

Dans les aubes mobiles, l'air de refroidissement pénètre dans les pales par le pied d'aube, mais dans les aubes fixes, l'air de refroidissement peut être introduit par une embase de l'aube fixe, soit en pied d'aube soit en tête d'aube, le pied d'aube étant l'extrémité de l'aube la plus proche de l'axe de rotation de la turbine.In the moving blades, the cooling air enters the blades by the blade root, but in the fixed blades, the air of cooling can be introduced by a fixed vane base, either at the foot of dawn is at the head of dawn, the foot of dawn being the end of the blade closest to the axis of rotation of the turbine.

On connait par US-A-4 992 026 une aube de turbine comportant une paroi aérodynamique creuse qui s'étend radialement entre un pied d'aube et une tête d'aube et qui présente un bord d'attaque et un bord de fuite, séparés l'un de l'autre et reliés par une paroi latérale concave ou intrados et une paroi latérale convexe ou extrados et comportant en outre un dispositif de refroidissement prévu à l'intérieur de ladite aube, alimenté en air de refroidissement par le pied d'aube et destiné à diriger l'air de refroidissement contre les surfaces intérieures desdites parois latérales.US-A-4 992 026 discloses a turbine blade comprising a hollow aerodynamic wall which extends radially between a foot blade and a blade head and which has a leading edge and an edge leakage, separated from each other and connected by a concave side wall or lower surface and a convex or upper side wall and comprising in addition to a cooling device provided inside said blade, supplied with cooling air from the blade root and intended for direct the cooling air against the interior surfaces of said side walls.

En outre, cette aube comporte deux cloisons radiales qui relient lesdites parois latérales concave et convexe et qui séparent l'intérieur de ladite aube en une cavité amont située près du bord d'attaque, une cavité médiane située entre lesdites cloisons radiales et une cavité aval située du côté du bord de fuite,
la cavité amont et la cavité médiane sont alimentées en air par une entrée prévue en pied d'aube, cet air s'évacuant ensuite desdites cavités par des orifices ménagés en tête d'aube, tandis que la cavité aval est alimentée en air par une entrée séparée prévue en pied d'aube, cet air s'évacuant ensuite par une pluralité de fentes ménagées dans le bord de fuite.In addition, this blade comprises two radial partitions which connect said concave and convex side walls and which separate the interior of said blade into an upstream cavity located near the leading edge, a middle cavity located between said radial partitions and a downstream cavity located on the trailing edge side,
the upstream cavity and the middle cavity are supplied with air by an inlet provided at the root of the blade, this air then being evacuated from said cavities by orifices provided at the head of the blade, while the downstream cavity is supplied with air by a separate inlet provided at the foot of the blade, this air then being evacuated by a plurality of slots formed in the trailing edge.

Le but de l'invention est de proposer une aube de turbine dans laquelle le dispositif de refroidissement utilise au mieux les capacités de l'air de refroidissement, afin de réduire le débit de ventilation et donc d'augmenter le rendement du moteur. Selon l'invention, cette aube est caractérisée par le fait que le dispositif de refroidissement comporte

  • dans la cavité amont, une rampe hélicoïdale qui s'étend entre le pied d'aube et la tête d'aube,
  • dans la cavité médiane, une chemise prenant appui sur les parois internes des cloisons radiales et maintenue à distance des parois latérales de l'aube par des éléments en saillie, cette chemise présentant en face des parois latérales de l'aube une pluralité d'orifices pour refroidir ces parois latérales par impact, et
  • dans la cavité aval, une cloison transversale obturant l'extrémité inférieure de ladite cavité et une troisième cloison radiale séparant ladite cavité en une partie amont et une partie aval près du bord de fuite, ces deux parties communiquant entre elles par une ouverture prévue en pied de ladite troisième cloison, et les cloisons latérales de l'aube au droit de la partie amont étant constituées de doubles peaux reliées par des pontets, et entre lesquelles circule un débit d'air de refroidissement introduit en pied d'aube, ce débit pénétrant ensuite dans la partie amont en tête d'aube, puis entrant dans la partie aval par ladite ouverture d'où elle s'évacue par la pluralité de fentes.
The object of the invention is to propose a turbine blade in which the cooling device makes the best use of the capacities of the cooling air, in order to reduce the ventilation rate and therefore to increase the efficiency of the engine. According to the invention, this blade is characterized in that the cooling device comprises
  • in the upstream cavity, a helical ramp which extends between the blade root and the blade head,
  • in the median cavity, a jacket bearing on the internal walls of the radial partitions and kept at a distance from the side walls of the blade by projecting elements, this jacket having a plurality of orifices opposite the side walls of the blade to cool these side walls by impact, and
  • in the downstream cavity, a transverse partition closing the lower end of said cavity and a third radial partition separating said cavity into an upstream part and a downstream part near the trailing edge, these two parts communicating with each other by an opening provided at the bottom of said third partition, and the side partitions of the blade in line with the upstream part being made up of double skins connected by bridges, and between which circulates a flow of cooling air introduced at the foot of the blade, this penetrating flow then in the upstream part at the head of the blade, then entering the downstream part through said opening from where it is evacuated by the plurality of slots.

Avantageusement la paroi interne de la cavité amont comporte des perturbateurs. Ces perturbateurs peuvent être constitués de nervures, de picots ou de pontets reliant la paroi interne de l'aube à l'âme de la rampe hélicoïdale.Advantageously, the internal wall of the upstream cavity comprises disruptors. These disruptors can consist of ribs, spikes or bridges connecting the inner wall of the blade to the soul of the helical ramp.

La chemise de la cavité médiane comporte avantageusement une pluralité de compartiments juxtaposés qui sont alimentés successivement par un même débit d'air. Le premier compartiment est alimenté en air par le pied d'aube, et les compartiments suivants sont alimentés par le débit d'air issu du compartiment précédent et ayant impacté les parois latérales de l'aube, par des fentes prévues dans les parois de la chemise sous les éléments en saillie, ces derniers étant constitués de nervures transversales. The jacket of the middle cavity advantageously comprises a plurality of juxtaposed compartments which are powered successively by the same air flow. The first compartment is supplied with air by the blade root, and the following compartments are supplied by the air flow coming from the previous compartment and having impacted the side walls of the dawn, by slots provided in the jacket walls under the protruding elements, the latter being made up of transverse ribs.

La rampe hélicoïdale permet d'augmenter très significativement le coefficient d'échange interne pour le refroidissement de l'aube dans la zone du bord d'attaque.The helical ramp makes it possible to increase very significantly the internal exchange coefficient for cooling the blade in the leading edge area.

Le système d'impact en cascade, disposé dans la cavité médiane, permet d'utiliser tout le potentiel de l'air de refroidissement avant qu'il ne soit réintroduit dans la veine.The cascade impact system, placed in the cavity middle, allows to use the full potential of the cooling air before it is reintroduced into the vein.

Avec le système à pontets prévu dans la cavité aval, on dispose d'un système de refroidissement efficace, proche des zones chaudes, et très facilement modulable.With the bridge system provided in the downstream cavity, we have an efficient cooling system, close to hot areas, and very easily modular.

La combinaison de ces technologies de refroidissement permet d'optimiser la ventilation des aubages de turbine en utilisant au maximum le potentiel de l'air de refroidissement et en ayant un dimensionnement thermique conduisant à une durée de vie mécanique optimale.The combination of these cooling technologies allows optimize the ventilation of the turbine blades by using at maximum the potential of the cooling air and having a thermal design leading to a mechanical service life optimal.

La conception de l'aube selon l'invention permet de réduire le débit de ventilation et donc d'augmenter le rendement du moteur.The design of the blade according to the invention makes it possible to reduce the ventilation flow and therefore increase the efficiency of the motor.

D'autres avantages et caractéristiques de l'invention ressortiront à la lecture de la description suivante faite à titre d'exemple non-limitatif et en référence aux dessins annexés dans lesquels :

  • la figure 1 est une vue de dessus de l'aube selon l'invention ;
  • la figure 2 est une coupe axiale de l'aube de la figure 1, cette coupe étant faite selon la surface axiale curviligne représentée par la ligne II-II sur la figure 1 ;
  • la figure 3 est une vue en perspective de la rampe hélicoïdale montée dans la cavité amont ;
  • les figures 4 à 7 sont des écorchés du bord d'attaque de l'aube, qui montrent la disposition de la rampe hélicoïdale dans la cavité amont, et divers types de perturbateurs ;
  • les figures 8 à 10 sont des coupes transversales de l'aube, prises à différentes distances du pied d'aube, respectivement selon les lignes VIII-VIII, IX-IX et X-X de la figure 2 ;
  • la figure 11 est une coupe de l'aube de la figure 2 faite selon un plan radial passant par un axe médian de la cavité médiane et représenté par la ligne XI-XI sur la figure 2 ;
  • la figure 12 est une coupe de l'aube de la figure 2 selon un plan radial coupant la cavité aval et représenté par la ligne XII-XII sur la figure 2 ;
  • la figure 13 est une coupe selon un plan médian d'une double peau formant la paroi externe de la cavité aval, plan représenté par la ligne XIII-XIII sur la figure 12 ;
  • la figure 14 est semblable à la figure 13 et montre une autre disposition des pontets reliant les doubles peaux.
    • Other advantages and characteristics of the invention will emerge on reading the following description given by way of non-limiting example and with reference to the appended drawings in which:
  • Figure 1 is a top view of the blade according to the invention;
  • Figure 2 is an axial section of the blade of Figure 1, this section being taken along the curvilinear axial surface shown by line II-II in Figure 1;
  • Figure 3 is a perspective view of the helical ramp mounted in the upstream cavity;
  • Figures 4 to 7 are cutaway of the leading edge of the blade, which show the arrangement of the helical ramp in the upstream cavity, and various types of disturbers;
  • Figures 8 to 10 are cross sections of the blade, taken at different distances from the blade root, respectively along the lines VIII-VIII, IX-IX and XX of Figure 2;
  • Figure 11 is a section of the blade of Figure 2 made along a radial plane passing through a median axis of the median cavity and represented by the line XI-XI in Figure 2;
  • Figure 12 is a section of the blade of Figure 2 along a radial plane intersecting the downstream cavity and shown by line XII-XII in Figure 2;
  • Figure 13 is a section along a median plane of a double skin forming the outer wall of the downstream cavity, plane represented by the line XIII-XIII in Figure 12;
  • Figure 14 is similar to Figure 13 and shows another arrangement of the bridges connecting the double skins.

    • Le dessin montre une aube mobile 1 d'une turbine à haute pression qui comporte une paroi aérodynamique creuse 2, également dénommée pale qui s'étend radialement entre un pied d'aube 3 et une tête d'aube 4. La paroi aérodynamique 2 présente quatre zones distinctes: un bord d'attaque 5 arrondi destiné à être disposé en regard du flux de gaz chauds issus de la chambre de combustion, un bord de fuite effilé 6, éloigné du bord d'attaque et relié à ce dernier par une paroi latérale concave 7, dénommée intrados, et une paroi latérale convexe 8, dénommée extrados.The drawing shows a moving blade 1 of a high turbine pressure which has a hollow aerodynamic wall 2, also referred to as a blade which extends radially between a blade root 3 and a blade head 4. The aerodynamic wall 2 has four zones distinct: a rounded leading edge 5 intended to be placed opposite the flow of hot gases from the combustion chamber, an edge of tapered leak 6, distant from the leading edge and connected to the latter by a concave side wall 7, called the lower surface, and a side wall convex 8, called upper surface.

    Les parois latérales 7 et 8 sont reliées par deux cloisons radiales 9 et 10 qui séparent l'intérieur de l'aube 1 en trois cavités : une cavité amont 11 située au voisinage immédiat du bord d'attaque 5, une cavité médiane 12 située entre les deux cloisons radiales 9 et 10 et une cavité aval 13 située du côté du bord de fuite 6. La cavité aval 13 est la plus large et occupe environ les deux tiers de l'étendue de l'aube 1.The side walls 7 and 8 are connected by two radial partitions 9 and 10 which separate the interior of the blade 1 into three cavities: a cavity upstream 11 located in the immediate vicinity of the leading edge 5, a cavity median 12 located between the two radial partitions 9 and 10 and a downstream cavity 13 located on the trailing edge side 6. The downstream cavity 13 is the widest and occupies about two-thirds of the extent of dawn 1.

    Une troisième cloison radiale 14 sépare en outre la cavité aval 13 en une partie amont 15 et une partie aval 16 près du bord de fuite 6. Une cloison transversale 17 obture l'extrémité inférieure de la cavité aval 13. La partie amont 15 et la partie aval 16 communiquent entre elles par une ouverture 18 ménagée en pied de la troisième cloison 14. Dans la partie effilée du bord de fuite 6 sont ménagées une pluralité de fentes 19 qui mettent en communication la partie aval 16 de la cavité aval 13 avec la veine de gaz de combustion qui s'écoule le long des parois latérales 7 et 8 de l'aube 1.A third radial partition 14 further separates the downstream cavity 13 in an upstream part 15 and a downstream part 16 near the trailing edge 6. A transverse partition 17 closes the lower end of the downstream cavity 13. The upstream part 15 and the downstream part 16 communicate between them by an opening 18 formed at the foot of the third partition 14. In the tapered part of the trailing edge 6 are provided a plurality of slots 19 which connect the downstream part 16 of the downstream cavity 13 with the combustion gas stream which flows along the side walls 7 and 8 of the blade 1.

    Ainsi qu'on le voit sur les figures 1 et 2, un orifice 20 est ménagé dans la paroi de la tête d'aube 4 au droit de la cavité amont 11, et un deuxième orifice 21, de forme oblongue, est ménagé dans la tête d'aube 4, au dessus de la cavité médiane 12. As seen in Figures 1 and 2, an orifice 20 is formed in the wall of the blade head 4 in line with the upstream cavity 11, and a second orifice 21, of oblong shape, is formed in the blade head 4, above the middle cavity 12.

    Dans le pied d'aube 3, sont ménagés deux conduits séparés 22 et 23 destinés à fournir de l'air de refroidissement. Le premier conduit 22 alimente directement en air de refroidissement les extrémités inférieures de la cavité amont 11 et de la cavité médiane 13, ainsi que cela est montré sur les figures 2 et 11, tandis que le deuxième conduit 23 alimente en air de refroidissement la partie amont 15 de la cavité aval 13 au voisinage de la tête d'aube 4, cet air ayant transité à l'intérieur des deux parois latérales 6 et 7 constituées de doubles peaux reliées par des pontets 24 au moins droit de la partie amont 15, ainsi que cela est représenté sur les figures 12 à 14.In the blade root 3, two separate conduits 22 are formed. and 23 for supplying cooling air. The first conduit 22 directly supplies cooling air to the ends lower of the upstream cavity 11 and the middle cavity 13, as well as this is shown in Figures 2 and 11, while the second leads 23 supplies cooling air to the upstream part 15 of the cavity downstream 13 in the vicinity of the blade head 4, this air having passed through the interior of the two side walls 6 and 7 consisting of double skins connected by at least straight bridges 24 of the upstream part 15, as shown in Figures 12 to 14.

    L'aube 1 est réalisée au niveau de sa paroi aérodynamique creuse 2 en deux demi-aubes réunies ultérieurement par brasage, la coupure entre les deux demi-aubes se faisant au niveau du squelette, ou l'aube peut être réalisée en fonderie.The blade 1 is made at its aerodynamic wall hollow 2 into two half-blades subsequently joined by brazing, the cut between the two half-blades at the skeleton, or the dawn can be made in foundry.

    Ainsi qu'on le voit sur les figures 2 à 7, la cavité amont 11 située près du bord d'attaque 5 est refroidie par convection par l'intermédiaire d'une rampe hélicoïdale 30.As seen in Figures 2 to 7, the upstream cavity 11 located near the leading edge 5 is cooled by convection by through a helical ramp 30.

    Cette rampe 30 peut être obtenue par fonderie et être monobloc avec une demi-aube, ou bien rapportée dans la cavité amont 11 et brasée.This ramp 30 can be obtained by foundry and be in one piece with a half-blade, or added in the upstream cavity 11 and brazed.

    Dans ce dernier cas, on a intérêt à utiliser un matériau avec une forte conductivité pour augmenter l'efficacité du refroidissement de ce circuit de ventilation.In the latter case, it is advantageous to use a material with a high conductivity to increase the cooling efficiency of this ventilation circuit.

    La rampe hélicoïdale 30 représentée sur la figure 3 comporte deux filets 31a, 31b, cependant cette rampe 30 peut ne posséder qu'un seul filet ou plus de deux, suivant les besoins.The helical ramp 30 shown in Figure 3 includes two nets 31a, 31b, however this ramp 30 may have only one single net or more than two, as needed.

    Le corps central 32, ou âme, de la rampe 30 n'est pas nécessairement cylindrique, il peut avoir une section évolutive sur la hauteur dans le but de moduler à souhait la section la section de passage de l'air de refroidissement afin de réguler le niveaux de coefficient d'échange.The central body 32, or core, of the ramp 30 is not necessarily cylindrical, it can have an evolving section on the height in order to modulate the section as desired the section of cooling air passage in order to regulate the levels of exchange coefficient.

    Dans la cavité amont 11, l'air de refroidissement circule dans un système de refroidissement du type "vis-sans-fin" qui part du pied 3 de l'aube et se termine à la tête de l'aube 5, d'où l'air s'évacue par l'orifice 20. Ce système permet d'augmenter sensiblement le parcours de l'écoulement d'air et d'augmenter, à débit de refroidissement fixé, la vitesse de l'écoulement par rapport à celle obtenue dans une cavité purement radiale.In the upstream cavity 11, the cooling air circulates in a "worm" type cooling system that starts from the bottom 3 of dawn and ends at the head of dawn 5, from which the air is evacuated by orifice 20. This system makes it possible to significantly increase the distance of the air flow and increase, at a fixed cooling rate, flow velocity relative to that obtained in a cavity purely radial.

    Le niveau de coefficient d'échange se trouve ainsi renforcé. De plus, cet écoulement tournant aura tendance à accentuer l'échange sur la paroi de l'aube au voisinage du bord d'attaque 5, l'air étant projeté sur l'extérieur de la rampe hélicoïdale 30 par effet centrifuge.The level of exchange coefficient is thus reinforced. Of more, this rotating flow will tend to accentuate the exchange on the wall of the blade in the vicinity of the leading edge 5, the air being projected on the outside of the helical ramp 30 by centrifugal effect.

    Comme on le voit sur les figures 4 à 7, plusieurs aménagements sont proposés en association avec la rampe hélicoïdale 30.As seen in Figures 4 to 7, several arrangements are available in combination with the helical ramp 30.

    Sur la figure 4, la rampe hélicoïdale est placée dans la cavité amont 11 dont la paroi interne est lisse.In Figure 4, the helical ramp is placed in the cavity upstream 11 whose internal wall is smooth.

    Sur la figure 5, on voit que des perturbateurs 33 sous la forme de nervures inclinées sont disposées soit sur la paroi interne de la cavité amont 11, soit sur la rampe hélicoïdale.In FIG. 5, it can be seen that disturbers 33 in the form inclined ribs are arranged either on the inner wall of the upstream cavity 11, ie on the helical ramp.

    Ainsi qu'on le voit sur la figure 6, les perturbateurs peuvent être constitués de pontets 34 qui relient la paroi interne de la cavité amont 11 à l'âme 32 de la rampe hélicoïdale 30. Ces pontets 34 peuvent être disposés en quinconce.As seen in Figure 6, the disruptors can be made up of bridges 34 which connect the internal wall of the upstream cavity 11 to the core 32 of the helical ramp 30. These bridges 34 can be staggered.

    Sur la figure 7, on voit que les perturbateurs peuvent être constitués par des picots 35 disposés en quinconce ou non sur la paroi interne de la cavité amont 11.In FIG. 7, it can be seen that the disturbers can be constituted by pins 35 arranged staggered or not on the wall internal of the upstream cavity 11.

    Le dispositif de refroidissement décrit ci-dessus est mis en place dans la cavité amont 11 situé au voisinage immédiat du bord d'attaque 5. Ce dispositif pourrait également être placé dans d'autres cavités.The cooling device described above is implemented place in the upstream cavity 11 located in the immediate vicinity of the edge 5. This device could also be placed in other rooms. cavities.

    L'air de refroidissement dans cette cavité amont 11 circule de manière centrifuge, du pied d'aube 3 vers la tête d'aube 5. Mais le circuit peut être inversé, notamment dans les aubes fixes des distributeurs de turbine, par exemple. Plusieurs rampes hélicoïdales peuvent également équiper une cavité avec retournement du circuit de refroidissement en pied ou en tête d'aube.The cooling air in this upstream cavity 11 circulates from centrifugal way, from the blade root 3 to the blade head 5. But the circuit can be reversed, especially in the fixed blades of turbine distributors, for example. Several helical ramps can also equip a cavity with reversal of the circuit cooling at the foot or at the head of the blade.

    La cavité médiane 12 est refroidie par convexion en utilisant la technologie d'impact en cascade par un air de refroidissement introduit dans la partie inférieure de la cavité 12 à partir du conduit 22 ménagé dans le pied d'aube 3.The central cavity 12 is cooled by convection using the cascade impact technology with cooling air introduced into the lower part of the cavity 12 from the conduit 22 formed in the blade root 3.

    Les figures 2 et 8 à 11 montrent qu'une chemise 40 est introduite dans la cavité médiane 12. Cette chemise 40 est réalisée par un assemblage mécano-soudé d'un ensemble de tôles préalablement percées pour réaliser des orifices d'impact 41, et des fentes 42 ou peut être réalisée directement en fonderie.Figures 2 and 8 to 11 show that a jacket 40 is introduced into the middle cavity 12. This jacket 40 is produced by a mechanically welded assembly of a set of sheets beforehand drilled to make impact holes 41, and slots 42 or can be carried out directly in the foundry.

    La chemise 40 se présente sous la forme d'une cheminée, dont deux parois latérales opposées 43 et 44 prennent appui sur les parois internes des cloisons radiales 9 et 10 et dont les deux autres parois opposées 45 et 46, qui comportent les orifices d'impact 41 et les fentes 42, sont maintenues à une certaine distance des parois latérales 7 et 8 de l'aube 1 par des éléments en saillie 47, sous forme de nervures transversales, formées sur les parois 45 et 46 et régulièrement répartis entre le pied d'aube 3 et la tête d'aube 4.The shirt 40 is in the form of a chimney, of which two opposite side walls 43 and 44 bear on the walls internal of the radial partitions 9 and 10 and of which the two other walls 45 and 46, which include the impact holes 41 and the slots 42 are kept at a certain distance from the side walls 7 and 8 of dawn 1 by projecting elements 47, in the form of transverse ribs, formed on the walls 45 and 46 and regularly distributed between the blade root 3 and the blade head 4.

    La cavité interne de la chemise 40 est partagée en un certain nombre de compartiments, référencés C1 à C7 sur la figure 11, au moyen de cloisons transversales 48 disposées respectivement, en partant du pied d'aube 3, sous un couple d'éléments en saillie 47 et séparées de ces éléments en saillie 47 par deux fentes 42 en regard des parois 7 et 8 de l'aube 1. La cloison supérieure 48a est écartée de la paroi formant la tête d'aube 4, afin que l'air de refroidissement évacué de la cavité C7 puisse s'évacuer par l'orifice 21.The internal cavity of the shirt 40 is divided into a certain number of compartments, referenced C1 to C7 in Figure 11, at by means of transverse partitions 48 arranged respectively, in starting from the blade root 3, under a couple of projecting elements 47 and separated from these projecting elements 47 by two facing slots 42 walls 7 and 8 of the blade 1. The upper partition 48a is separated from the wall forming the blade head 4, so that the cooling air evacuated from the cavity C7 can be evacuated through the orifice 21.

    Le circuit de refroidissement dans la cavité médiane 12 s'effectue de la manière suivanteThe cooling circuit in the middle cavity 12 is carried out as follows

    L'air est amené par le conduit 22 dans le compartiment C1 de la chemise 40, puis est évacué du compartiment C1 par les orifices d'impact 41, afin de frapper les parois internes de l'intrados 7 et de l'extrados 8 de l'aube 1 au voisinage du pied d'aube 3. Après impact, l'air est introduit dans le deuxième compartiment C2 par les premières fentes 42, puis évacué par les orifices d'impact 21 du compartiment C2 pour être ensuite réintroduit dans le troisième compartiment C3. L'air circule de cette manière jusqu'au compartiment supérieur C7, d'où il impacte les parois internes de l'intrados 7 et de l'extrados 8 au voisinage de la tête d'aube 4, puis est évacué hors de l'aube L par l'orifice 21.The air is brought through line 22 into compartment C1 of the jacket 40, then is evacuated from compartment C1 through the orifices impact 41, in order to strike the internal walls of the lower surface 7 and the upper surface 8 of the blade 1 in the vicinity of the blade root 3. After impact, air is introduced into the second compartment C2 through the first slots 42, then discharged through the impact orifices 21 of the compartment C2 to be then reintroduced into the third compartment C3. The air flows in this way to the upper compartment C7, from where it impacts the internal walls of the lower surface 7 and the upper surface 8 neighborhood of the blade head 4, then is evacuated out of the blade L by orifice 21.

    Le nombre de compartiments peut être différent de 7, et le nombre d'orifices d'impact 41 peut être différent d'un compartiment à l'autre. The number of compartments can be different from 7, and the number of impact orifices 41 may be different from compartment to the other.

    La chemise 40 décrite ci-dessus pourrait également être montée dans une cavité voisine du bord d'attaque ou du bord de fuite. Elle peut s'adapter aussi bien aux aubages fixes qu'aux aubages mobiles. Pour les aubages fixes, l'alimentation peut se faire par la tête d'aube 4, et les compartiments C1 à C7 peuvent être disposés radialement, comme dans l'exemple décrit ci-dessus, ou être disposés axialement du bord d'attaque 5 vers le bord de fuite 6 ou inversement. Ce dispositif peut s'appliquer aussi bien pour de l'impact réparti (plusieurs rangées d'orifices) que pour de l'impact concentré (une seule rangée d'orifices 41).The shirt 40 described above could also be mounted in a cavity near the leading edge or the trailing edge. She can be adapted to both fixed and mobile blades. For fixed blades, the feed can be done by the blade head 4, and compartments C1 to C7 can be arranged radially, as in the example described above, or be arranged axially from the leading edge 5 to the trailing edge 6 or vice versa. This device can be applied as well for distributed impact (several rows of orifices) only for concentrated impact (a single row of holes 41).

    Ainsi que cela a été mentionné plus haut, l'intrados 7 et l'extrados 8 comportent au niveau de la partie amont 15 de la cavité aval 13 des double peaux 7a, 7b et 8a, 8b, reliées par des pontets 24. Les peaux internes 7b, 8b sont reliées au voisinage du pied d'aube 3 par la cloison transversale 17. Ces deux peaux internes 7b, 8b s'étendent jusqu'au voisinage de la cloison formant la tête d'aube 4, tout en réservant des passages 50a, 50b près de la tête d'aube 4 par lesquels, l'air introduit par l'orifice 23 du pied d'aube 3, et ayant circulé de manière centrifuge entre les peaux 7a, 7b de l'intrados 7 et les peaux 8a, 8b de l'extrados 8, s'évacue dans la partie amont 15 de la cavité aval. Cet air de refroidissement circule de manière centripète dans cette partie amont 15, puis entre dans la partie aval 16 par l'ouverture 18. L'air remonte enfin de manière centrifuge dans la partie aval 16 et s'évacue dans la veine de gaz chauds par les fentes 19 ménagées dans le bord de fuite 6. L'air de refroidissement introduit par l'orifice 23 est divisé en deux débits B1 et B2 par la cloison transversale 17. Ces deux débits B1 et B2 circulent de manière centrifuge au travers de la multitude de pontets 24. Ces pontets 24 sont obtenus en fonderie lors de la coulée. Ces pontets 24 peuvent être disposés en quinconce (voir figure 13) ou disposés en ligne (voir figure 14). La forme des pontets peut être quelconque, de section cylindrique, carrée, oblongue .... Ce dispositif peut également être utilisé pour le refroidissement des zones s'étendant jusqu'au bord d'attaque.As mentioned above, the lower surface 7 and the upper surface 8 comprise at the level of the upstream part 15 of the cavity downstream 13 of the double skins 7a, 7b and 8a, 8b, connected by bridges 24. The internal skins 7b, 8b are connected in the vicinity of the blade root 3 by the transverse partition 17. These two internal skins 7b, 8b extend to the vicinity of the partition forming the blade head 4, while reserving passages 50a, 50b near the blade head 4 by which, the air introduced through the orifice 23 of the blade root 3, and having circulated centrifugally between the skins 7a, 7b of the lower surface 7 and the skins 8a, 8b of the upper surface 8, is evacuated in the upstream part 15 of the downstream cavity. This cooling air circulates centripetally in this upstream part 15, then enters the downstream part 16 by the opening 18. The air finally rises centrifugally in the downstream part 16 and is discharged into the stream of hot gases through the slots 19 formed in the trailing edge 6. The cooling air introduced by orifice 23 is divided into two flows B1 and B2 by the partition transverse 17. These two flows B1 and B2 circulate so centrifugal through the multitude of bridges 24. These bridges 24 are obtained in foundry during casting. These bridges 24 can be staggered (see Figure 13) or arranged in a row (see figure 14). The shape of the bridges can be any, of section cylindrical, square, oblong .... This device can also be used for cooling areas extending to the edge attack.

    La constitution des circuits internes de refroidissement se réalise en assemblant les pièces rapportées, rampe hélicoïdale 30 et chemise 40 mécano-soudée, dans une des demi-aubes, puis en rapportant l'autre demi-aube sur la précédente et en brasant ensuite l'ensemble des pièces. Les circuits de refroidissement peuvent également être réalisés entièrement ou partiellement directement en fonderie.The constitution of the internal cooling circuits is realizes by assembling the added parts, helical ramp 30 and 40 welded shirt, in one of the half-blades, then in bringing the other half-dawn over the previous one and then brazing all the parts. The cooling circuits can also be carried out entirely or partially directly in foundry.

    Claims (7)

    1. Turbine blade comprising a hollow aerodynamic wall (2) extending radially between a blade root (3) and a blade tip (4) and having a leading edge (5) and a trailing edge (6) separated from one another and connected by a concave lateral wall (7) or pressure face and a convex lateral wall (8) or suction face, further comprising a cooling device provided inside the said blade, supplied with cooling air via the blade root (3) and intended to direct the cooling air against the interior surfaces of the said lateral walls, and comprising two radial partitions (9, 10) connecting the said concave (7) and convex (8) lateral walls and dividing the inside of the said blade (1) into an upstream cavity (11) situated near the leading edge (5), a middle cavity (12) situated between the said radial partitions (9, 10) and a downstream cavity (13) situated at the trailing edge (6) end, the upstream cavity (11) and the middle cavity (12) being supplied with air via an inlet (22) made in the blade root (3), this air then being removed from the said cavities (11, 12) by orifices (20, 21) made in the blade tip (4), while the downstream cavity (13) is supplied with air by a separate inlet (23) provided at the blade root (3), this air then being removed through a plurality of slots (19) formed in the trailing edge (6), characterized in that the cooling device comprises:
      in the upstream cavity (11), a helical ramp (30) which stretches between the blade root (3) and the blade tip (4),
      in the middle cavity (12), a liner (40) resting against the internal walls of the radial partitions (9, 10) and held away from the lateral walls (7, 8) of the blade (1) by projecting elements (47), this liner (40) having, facing the lateral walls (7, 8) of the blade, a plurality of orifices (41) for cooling these lateral walls (7, 8) by impact, and
      in the downstream cavity (13), a transverse partition (17) closing off the lower end of the said cavity (13) and a radial third partition (14) dividing the said cavity (13) into an upstream part (15) and a downstream part (16) near the trailing edge (6), these two parts (15, 16) communicating with one another through an opening (18) made at the foot of the said third partition (14), and the lateral partitions (7, 8) of the blade at the upstream part (15) consisting of double shell-walls (7a, 7b; 8a, 8b) connected by ribs (24) and between which cooling air introduced at the blade root (3) flows, this flow then entering the upstream part (15) at the blade tip (4), then entering the downstream part (16) through the said opening (18) from where it is removed through the plurality of slots (19).
    2. Blade according to Claim 1, characterized in that the internal wall of the upstream cavity (13) has disrupters (33, 34, 35).
    3. Blade according to Claim 2, characterized in that the disrupters consist of ridges (33).
    4. Blade according to Claim 2, characterized in that the disrupters consist of ribs (34) connecting the internal wall of the blade to the core (32) of the helical ramp.
    5. Blade according to Claim 2, characterized in that the disrupters consist of spikes (35).
    6. Blade according to one of Claims 1 to 5,
      characterized in that the liner (40) of the middle cavity (13) has a plurality of juxtaposed compartments (C1 to C7) supplied in turn with the same air flow from the blade root (3).
    7. Blade according to Claim 6, characterized in that the compartments (C2 to C7), except for the first one, are supplied with the air flow from the previous compartment (C1 to C6), which flow has impacted on the lateral walls (7, 8) of the blade, via slots (42) made in the walls (45, 46) of the liner (40) under the projecting elements (47), the latter elements consisting of transverse ridges.
    EP98401558A 1997-06-26 1998-06-25 Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall Expired - Lifetime EP0887515B1 (en)

    Applications Claiming Priority (2)

    Application Number Priority Date Filing Date Title
    FR9707988A FR2765265B1 (en) 1997-06-26 1997-06-26 BLADED COOLING BY HELICAL RAMP, CASCADE IMPACT AND BY BRIDGE SYSTEM IN A DOUBLE SKIN
    FR9707988 1997-06-26

    Publications (2)

    Publication Number Publication Date
    EP0887515A1 EP0887515A1 (en) 1998-12-30
    EP0887515B1 true EP0887515B1 (en) 2003-08-13

    Family

    ID=9508460

    Family Applications (1)

    Application Number Title Priority Date Filing Date
    EP98401558A Expired - Lifetime EP0887515B1 (en) 1997-06-26 1998-06-25 Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall

    Country Status (6)

    Country Link
    US (1) US5993156A (en)
    EP (1) EP0887515B1 (en)
    JP (1) JP3735201B2 (en)
    DE (1) DE69817094T2 (en)
    FR (1) FR2765265B1 (en)
    RU (1) RU2146766C1 (en)

    Families Citing this family (77)

    * Cited by examiner, † Cited by third party
    Publication number Priority date Publication date Assignee Title
    GB2345942B (en) * 1998-12-24 2002-08-07 Rolls Royce Plc Gas turbine engine internal air system
    US6206638B1 (en) * 1999-02-12 2001-03-27 General Electric Company Low cost airfoil cooling circuit with sidewall impingement cooling chambers
    US6402470B1 (en) * 1999-10-05 2002-06-11 United Technologies Corporation Method and apparatus for cooling a wall within a gas turbine engine
    US6435814B1 (en) * 2000-05-16 2002-08-20 General Electric Company Film cooling air pocket in a closed loop cooled airfoil
    US6508627B2 (en) 2001-05-30 2003-01-21 Lau Industries, Inc. Airfoil blade and method for its manufacture
    US6609891B2 (en) * 2001-08-30 2003-08-26 General Electric Company Turbine airfoil for gas turbine engine
    US6981846B2 (en) 2003-03-12 2006-01-03 Florida Turbine Technologies, Inc. Vortex cooling of turbine blades
    US6932573B2 (en) 2003-04-30 2005-08-23 Siemens Westinghouse Power Corporation Turbine blade having a vortex forming cooling system for a trailing edge
    US7343232B2 (en) 2003-06-20 2008-03-11 Geneva Aerospace Vehicle control system including related methods and components
    FR2858352B1 (en) * 2003-08-01 2006-01-20 Snecma Moteurs COOLING CIRCUIT FOR TURBINE BLADE
    US6955525B2 (en) * 2003-08-08 2005-10-18 Siemens Westinghouse Power Corporation Cooling system for an outer wall of a turbine blade
    US7818127B1 (en) * 2004-06-18 2010-10-19 Geneva Aerospace, Inc. Collision avoidance for vehicle control systems
    DE502004008210D1 (en) * 2004-07-26 2008-11-20 Siemens Ag Cooled component of a turbomachine and method for casting this cooled component
    GB0418914D0 (en) * 2004-08-25 2004-09-29 Rolls Royce Plc Turbine component
    EP1655451B1 (en) * 2004-11-09 2010-06-30 Rolls-Royce Plc A cooling arrangement
    US7163373B2 (en) * 2005-02-02 2007-01-16 Siemens Power Generation, Inc. Vortex dissipation device for a cooling system within a turbine blade of a turbine engine
    RU2425982C2 (en) * 2005-04-14 2011-08-10 Альстом Текнолоджи Лтд Gas turbine vane
    US7563072B1 (en) * 2006-09-25 2009-07-21 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall spiral flow cooling circuit
    US7641445B1 (en) 2006-12-01 2010-01-05 Florida Turbine Technologies, Inc. Large tapered rotor blade with near wall cooling
    US7753650B1 (en) 2006-12-20 2010-07-13 Florida Turbine Technologies, Inc. Thin turbine rotor blade with sinusoidal flow cooling channels
    US7665965B1 (en) * 2007-01-17 2010-02-23 Florida Turbine Technologies, Inc. Turbine rotor disk with dirt particle separator
    US7901182B2 (en) * 2007-05-18 2011-03-08 Siemens Energy, Inc. Near wall cooling for a highly tapered turbine blade
    US20090060714A1 (en) * 2007-08-30 2009-03-05 General Electric Company Multi-part cast turbine engine component having an internal cooling channel and method of forming a multi-part cast turbine engine component
    FR2924156B1 (en) * 2007-11-26 2014-02-14 Snecma TURBINE DAWN
    US9322285B2 (en) * 2008-02-20 2016-04-26 United Technologies Corporation Large fillet airfoil with fanned cooling hole array
    US8297927B1 (en) * 2008-03-04 2012-10-30 Florida Turbine Technologies, Inc. Near wall multiple impingement serpentine flow cooled airfoil
    GB2462087A (en) * 2008-07-22 2010-01-27 Rolls Royce Plc An aerofoil comprising a partition web with a chordwise or spanwise variation
    US8303252B2 (en) * 2008-10-16 2012-11-06 United Technologies Corporation Airfoil with cooling passage providing variable heat transfer rate
    US8096766B1 (en) 2009-01-09 2012-01-17 Florida Turbine Technologies, Inc. Air cooled turbine airfoil with sequential cooling
    US8342797B2 (en) * 2009-08-31 2013-01-01 Rolls-Royce North American Technologies Inc. Cooled gas turbine engine airflow member
    US9528382B2 (en) * 2009-11-10 2016-12-27 General Electric Company Airfoil heat shield
    US8511994B2 (en) * 2009-11-23 2013-08-20 United Technologies Corporation Serpentine cored airfoil with body microcircuits
    US20120076660A1 (en) * 2010-09-28 2012-03-29 Spangler Brandon W Conduction pedestals for a gas turbine engine airfoil
    US9011077B2 (en) 2011-04-20 2015-04-21 Siemens Energy, Inc. Cooled airfoil in a turbine engine
    GB2498551B (en) * 2012-01-20 2015-07-08 Rolls Royce Plc Aerofoil cooling
    DE102012017491A1 (en) * 2012-09-04 2014-03-06 Rolls-Royce Deutschland Ltd & Co Kg Turbine blade of a gas turbine with swirl-generating element
    KR101317443B1 (en) * 2012-10-10 2013-10-10 한국항공대학교산학협력단 A cooled blade of gas turbine
    WO2014175951A2 (en) * 2013-03-15 2014-10-30 United Technologies Corporation Gas turbine engine component with twisted internal channel
    US10240470B2 (en) 2013-08-30 2019-03-26 United Technologies Corporation Baffle for gas turbine engine vane
    US10487668B2 (en) 2013-09-06 2019-11-26 United Technologies Corporation Gas turbine engine airfoil with wishbone baffle cooling scheme
    US20160222793A1 (en) * 2013-09-09 2016-08-04 United Technologies Corporation Cooling configuration for engine component
    EP2863010A1 (en) * 2013-10-21 2015-04-22 Siemens Aktiengesellschaft Turbine blade
    US8864438B1 (en) * 2013-12-05 2014-10-21 Siemens Energy, Inc. Flow control insert in cooling passage for turbine vane
    WO2015094531A1 (en) 2013-12-20 2015-06-25 United Technologies Corporation Gas turbine engine component cooling cavity with vortex promoting features
    KR101509385B1 (en) * 2014-01-16 2015-04-07 두산중공업 주식회사 Turbine blade having swirling cooling channel and method for cooling the same
    US20150204197A1 (en) * 2014-01-23 2015-07-23 Siemens Aktiengesellschaft Airfoil leading edge chamber cooling with angled impingement
    RU2568763C2 (en) * 2014-01-30 2015-11-20 Альстом Текнолоджи Лтд Gas turbine component
    WO2015160404A2 (en) * 2014-02-13 2015-10-22 United Technologies Corporation Gas turbine engine component with separation rib for cooling passages
    US10012090B2 (en) * 2014-07-25 2018-07-03 United Technologies Corporation Airfoil cooling apparatus
    FR3032173B1 (en) 2015-01-29 2018-07-27 Safran Aircraft Engines Blower blade of a blowing machine
    US10190420B2 (en) * 2015-02-10 2019-01-29 United Technologies Corporation Flared crossovers for airfoils
    US9849510B2 (en) 2015-04-16 2017-12-26 General Electric Company Article and method of forming an article
    US9915151B2 (en) * 2015-05-26 2018-03-13 Rolls-Royce Corporation CMC airfoil with cooling channels
    US9976441B2 (en) 2015-05-29 2018-05-22 General Electric Company Article, component, and method of forming an article
    US10087776B2 (en) 2015-09-08 2018-10-02 General Electric Company Article and method of forming an article
    US10739087B2 (en) * 2015-09-08 2020-08-11 General Electric Company Article, component, and method of forming an article
    US10253986B2 (en) 2015-09-08 2019-04-09 General Electric Company Article and method of forming an article
    RU2706211C2 (en) 2016-01-25 2019-11-14 Ансалдо Энерджиа Свитзерлэнд Аг Cooled wall of turbine component and cooling method of this wall
    EP3228819B1 (en) * 2016-04-08 2021-06-09 Ansaldo Energia Switzerland AG Blade comprising cmc layers
    US10156146B2 (en) * 2016-04-25 2018-12-18 General Electric Company Airfoil with variable slot decoupling
    FR3052183B1 (en) * 2016-06-02 2020-03-06 Safran Aircraft Engines TURBINE BLADE COMPRISING A COOLING AIR INTAKE PORTION INCLUDING A HELICOIDAL ELEMENT FOR SWIRLING THE COOLING AIR
    RU171631U1 (en) * 2016-09-14 2017-06-07 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения им. П.И. Баранова" Cooled turbine blade
    DE102016221009A1 (en) 2016-10-26 2018-04-26 Continental Reifen Deutschland Gmbh Pressure control device
    US20180149028A1 (en) * 2016-11-30 2018-05-31 General Electric Company Impingement insert for a gas turbine engine
    CN106703899B (en) * 2017-01-23 2019-08-23 中国航发沈阳发动机研究所 High Pressure Turbine Rotor blade inlet edge impinging cooling structure and the engine with it
    US10494948B2 (en) * 2017-05-09 2019-12-03 General Electric Company Impingement insert
    US10570751B2 (en) 2017-11-22 2020-02-25 General Electric Company Turbine engine airfoil assembly
    US10787912B2 (en) * 2018-04-25 2020-09-29 Raytheon Technologies Corporation Spiral cavities for gas turbine engine components
    US10787913B2 (en) * 2018-11-01 2020-09-29 United Technologies Corporation Airfoil cooling circuit
    US11149550B2 (en) * 2019-02-07 2021-10-19 Raytheon Technologies Corporation Blade neck transition
    US10871074B2 (en) 2019-02-28 2020-12-22 Raytheon Technologies Corporation Blade/vane cooling passages
    FR3107919B1 (en) 2020-03-03 2022-12-02 Safran Aircraft Engines Hollow turbomachine blade and inter-blade platform fitted with projections that disrupt cooling flow
    FR3108145B1 (en) * 2020-03-13 2022-02-18 Safran Helicopter Engines HOLLOW DAWN OF TURBOMACHINE
    CN112610284A (en) * 2020-12-17 2021-04-06 东北电力大学 Gas turbine blade with spiral band
    CN113374536B (en) * 2021-06-09 2022-08-09 中国航发湖南动力机械研究所 Gas turbine guide vane
    US20230417146A1 (en) * 2022-06-23 2023-12-28 Solar Turbines Incorporated Pneumatically variable turbine nozzle
    CN116950723B (en) * 2023-09-19 2024-01-09 中国航发四川燃气涡轮研究院 Low-stress double-wall turbine guide vane cooling structure and design method thereof

    Family Cites Families (16)

    * Cited by examiner, † Cited by third party
    Publication number Priority date Publication date Assignee Title
    DE853534C (en) * 1943-02-27 1952-10-27 Maschf Augsburg Nuernberg Ag Air-cooled gas turbine blade
    BE485284A (en) * 1947-10-28
    NL73916C (en) * 1949-07-06 1900-01-01
    DE2514208A1 (en) * 1975-04-01 1976-10-14 Kraftwerk Union Ag DISC DESIGN GAS TURBINE
    CH584833A5 (en) * 1975-05-16 1977-02-15 Bbc Brown Boveri & Cie
    US4173120A (en) * 1977-09-09 1979-11-06 International Harvester Company Turbine nozzle and rotor cooling systems
    US4407632A (en) * 1981-06-26 1983-10-04 United Technologies Corporation Airfoil pedestaled trailing edge region cooling configuration
    DE3306894A1 (en) * 1983-02-26 1984-08-30 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Turbine stator or rotor blade with cooling channel
    JPS62228603A (en) * 1986-03-31 1987-10-07 Toshiba Corp Gas turbine blade
    US5002460A (en) * 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade
    FR2678318B1 (en) * 1991-06-25 1993-09-10 Snecma COOLED VANE OF TURBINE DISTRIBUTOR.
    JP3006174B2 (en) * 1991-07-04 2000-02-07 株式会社日立製作所 Member having a cooling passage inside
    US5259730A (en) * 1991-11-04 1993-11-09 General Electric Company Impingement cooled airfoil with bonding foil insert
    US5695322A (en) * 1991-12-17 1997-12-09 General Electric Company Turbine blade having restart turbulators
    US5464322A (en) * 1994-08-23 1995-11-07 General Electric Company Cooling circuit for turbine stator vane trailing edge
    US5472316A (en) * 1994-09-19 1995-12-05 General Electric Company Enhanced cooling apparatus for gas turbine engine airfoils

    Also Published As

    Publication number Publication date
    JPH1172003A (en) 1999-03-16
    DE69817094D1 (en) 2003-09-18
    RU2146766C1 (en) 2000-03-20
    EP0887515A1 (en) 1998-12-30
    FR2765265B1 (en) 1999-08-20
    JP3735201B2 (en) 2006-01-18
    FR2765265A1 (en) 1998-12-31
    DE69817094T2 (en) 2004-06-17
    US5993156A (en) 1999-11-30

    Similar Documents

    Publication Publication Date Title
    EP0887515B1 (en) Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall
    CA2193165C (en) Cooled turbine vane
    CA2475083C (en) Cooling circuits for gas turbine blades
    EP0821201B1 (en) Screen-deflector assembly for the combustion chamber of a gas turbine
    EP0666406B1 (en) Cooled turbine blade
    CA2398659C (en) Cooling circuits for gas turbine blade
    CA2946708C (en) Turbomachine turbine blade comprising a cooling circuit with improved homogeneity
    EP1586743B1 (en) Turbine shroud
    FR2966868A1 (en) SYSTEM AND METHOD FOR COOLING ROTARY BLADE PLATFORM AREAS OF TURBINE
    FR2695162A1 (en) Fin with advanced end cooling system.
    EP0250323A1 (en) Control device for the flux of cooling air for the rotor of a turbine engine
    FR2887287A1 (en) Turbomachine e.g. high pressure turbine, rotor blade, has intrados and extrados cooling circuits with intrados and extrados cavities extending to central wall, and outlet orifices opened in central cavities and leading on intrados side
    FR2966869A1 (en) SYSTEM FOR COOLING TURBINE ROTARY BLADE PLATFORM ZONES
    EP1790819A1 (en) Cooling circuit for a turbine blade
    FR3021699A1 (en) OPTIMIZED COOLING TURBINE BLADE AT ITS LEFT EDGE
    FR2969210A1 (en) SYSTEM FOR COOLING ZONES OF ROTORIC TURBINE BLADE PLATFORMS
    FR2981979A1 (en) TURBINE WHEEL FOR A TURBOMACHINE
    EP3149281B1 (en) Turbine blade having a central cooling conduit and two lateral cavities merged downstream of the central conduit
    EP1630351B1 (en) Blade for a compressor or a gas turbine
    FR3028576A1 (en) TURBOMACHINE STATOR AUBING SECTOR COMPRISING HOT FLUID CIRCULATION CHANNELS
    FR3028575A1 (en) STATOR AUBING SECTOR OF A TURBOMACHINE
    CA3059400A1 (en) Blade comprising an improved cooling circuit
    EP3942157A1 (en) Turbine engine vane equipped with a cooling circuit and lost-wax method for manufacturing such a vane
    WO2020224995A1 (en) Turbine vane of a turbomachine, turbine, turbomachine and associated ceramic core for manufacturing a turbine vane of a turbomachine
    EP3867499A1 (en) Turbine engine blade with improved cooling

    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: 19980714

    AK Designated contracting states

    Kind code of ref document: A1

    Designated state(s): DE FR GB

    AX Request for extension of the european patent

    Free format text: AL;LT;LV;MK;RO;SI

    AKX Designation fees paid

    Free format text: DE FR GB

    17Q First examination report despatched

    Effective date: 20020711

    GRAH Despatch of communication of intention to grant a patent

    Free format text: ORIGINAL CODE: EPIDOS IGRA

    GRAH Despatch of communication of intention to grant a patent

    Free format text: ORIGINAL CODE: EPIDOS IGRA

    GRAA (expected) grant

    Free format text: ORIGINAL CODE: 0009210

    RAP1 Party data changed (applicant data changed or rights of an application transferred)

    Owner name: SNECMA MOTEURS

    AK Designated contracting states

    Designated state(s): DE FR GB

    REG Reference to a national code

    Ref country code: GB

    Ref legal event code: FG4D

    Free format text: NOT ENGLISH

    REF Corresponds to:

    Ref document number: 69817094

    Country of ref document: DE

    Date of ref document: 20030918

    Kind code of ref document: P

    GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)
    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: 20040514

    REG Reference to a national code

    Ref country code: FR

    Ref legal event code: CD

    REG Reference to a national code

    Ref country code: FR

    Ref legal event code: PLFP

    Year of fee payment: 19

    REG Reference to a national code

    Ref country code: FR

    Ref legal event code: PLFP

    Year of fee payment: 20

    PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

    Ref country code: DE

    Payment date: 20170522

    Year of fee payment: 20

    Ref country code: GB

    Payment date: 20170526

    Year of fee payment: 20

    Ref country code: FR

    Payment date: 20170427

    Year of fee payment: 20

    REG Reference to a national code

    Ref country code: FR

    Ref legal event code: CD

    Owner name: SAFRAN AIRCRAFT ENGINES

    Effective date: 20170713

    REG Reference to a national code

    Ref country code: DE

    Ref legal event code: R071

    Ref document number: 69817094

    Country of ref document: DE

    REG Reference to a national code

    Ref country code: GB

    Ref legal event code: PE20

    Expiry date: 20180624

    PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

    Ref country code: GB

    Free format text: LAPSE BECAUSE OF EXPIRATION OF PROTECTION

    Effective date: 20180624