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EP3658751B1 - Aubage d'aube de turbine - Google Patents

Aubage d'aube de turbine Download PDF

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Publication number
EP3658751B1
EP3658751B1 EP18779293.2A EP18779293A EP3658751B1 EP 3658751 B1 EP3658751 B1 EP 3658751B1 EP 18779293 A EP18779293 A EP 18779293A EP 3658751 B1 EP3658751 B1 EP 3658751B1
Authority
EP
European Patent Office
Prior art keywords
impingement cooling
blade
wall
cooling
impingement
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
Application number
EP18779293.2A
Other languages
German (de)
English (en)
Other versions
EP3658751A1 (fr
Inventor
Heinz-Jürgen GROSS
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.)
Siemens Energy Global GmbH and Co KG
Original Assignee
Siemens Energy Global GmbH and Co KG
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 Siemens Energy Global GmbH and Co KG filed Critical Siemens Energy Global GmbH and Co KG
Publication of EP3658751A1 publication Critical patent/EP3658751A1/fr
Application granted granted Critical
Publication of EP3658751B1 publication Critical patent/EP3658751B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/186—Film 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
    • 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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • 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
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/185—Two-dimensional patterned serpentine-like
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the invention relates to an airfoil for a turbine blade according to the preamble of claim 1, for example in FIG DE 10 2016 123 525 A1 disclosed.
  • An airfoil corresponding to the preamble of claim 1 has long been known from the extensive prior art.
  • the airfoil and in particular also the entire gas turbine blade are usually produced in an investment casting process, so that cavities are present in the interior of the airfoil.
  • a coolant mostly cooling air, can flow through these cavities so that the metallic material of the blade and the turbine blade can permanently withstand the high temperatures occurring during operation.
  • impingement cooling Different cooling concepts, which have long been known, are used for cooling, one of which is referred to as impingement cooling.
  • cooling air jets hit the inner surfaces of the metallic blade wall at an almost perpendicular angle in order to absorb the thermal energy contained therein and then to transport it away with them.
  • the impingement cooling walls required to form the impingement cooling can on the one hand be cast with or on the other hand be provided by installing metal sheet metal inserts.
  • cast impingement cooling systems require a minimum distance between the wall surface to be cooled and the impingement cooling wall having the impingement cooling openings, since the cast cores themselves required a minimum wall thickness for sufficient strength.
  • the perforated impact cooling wall is mounted as an insert in a blade, additional manufacturing and assembly steps are required, which increase the cost of manufacturing the turbine blade.
  • additional manufacturing and assembly steps are required, which increase the cost of manufacturing the turbine blade.
  • the object of the invention is therefore to provide a long-lasting airfoil for a turbine blade which enables particularly efficient cooling of the side walls of the airfoil.
  • the present invention proposes that in an airfoil for a turbine blade, comprising a suction-side side wall and a pressure-side side wall, which extend along a profile center line from a common leading edge to a common trailing edge and in a spanwise direction from a root-side end to a head-side end At least partially enclose the cavity, with a first perforated impact cooling wall for impingement cooling of the front edge and at least one further perforated impingement cooling wall for impingement cooling of a section of the suction-side and / or pressure-side side wall, the impingement cooling openings of the first impingement cooling wall and the impingement cooling openings of the at least one being provided along the span inside further baffle cooling wall are fluidically connected in series.
  • cascaded impingement cooling inside the blade is proposed, with at least one further impingement cooling section, preferably two further impingement cooling sections, per side wall cascading downstream from a first impingement cooling at the leading edge on the suction side and / or pressure side.
  • the invention is based on the knowledge that series-connected impingement cooling (cascaded impingement cooling) allows the cooling air to be used multiple times and thus to make it more uniform to achieve the temperature distribution along the cross-section.
  • the area of the airfoil which is thermally most heavily loaded, ie the area around the leading edge, is fed with the coolest cooling air and is impingely cooled in a first impingement cooling section.
  • the cooling air is heated up for the first time and the blade temperature in the vicinity of the leading edge is reduced to a tolerable level.
  • the heated cooling air is then passed into a downstream section of the airfoil and used there again for impingement cooling of the side wall, whereby the temperature of the side wall there is also lowered and the cooling air is heated again.
  • This achieves an efficient use of cooling air, so that - compared with conventional blades - the saved cooling air can be used to increase the efficiency of the gas turbine.
  • the thermal displacement across the blade cross-section can be reduced. This can reduce the thermo-mechanical load on the metallic airfoil, which can lead to an increased service life of the airfoil. Due to the fact that the series-connected impingement cooling has low cross-flow components in the spanwise direction, it is comparatively efficient.
  • An impingement cooling space is also provided between the relevant impingement cooling wall and the inside of the associated side wall, a collecting space being provided downstream of the impinging cooling space in question, which is immediately adjacent to the downstream further impingement cooling wall.
  • the relative terms “upstream” and “downstream” refer to the direction of flow of the cooling air inside the airfoil, unless otherwise stated.
  • the collecting spaces serve as cavities in which the coolant, which has been further heated after impingement cooling, can be collected on the one hand and it can be collected from the other through the impingement cooling openings of the subsequent ones Impact cooling wall can pass through for further impingement cooling.
  • the collecting spaces preferably extend in the spanwise direction over the entire length of the blade. As a result, the pressure in the collecting space can be made uniform.
  • the collecting space is partially delimited by a projection which is impact-cooled.
  • outlet openings close to the side wall are arranged in a rib, which extends according to a cross-sectional plane from a rib end on the suction side to a rib end on the pressure side.
  • a supply channel for supplying coolant for cooling the leading edge is provided between the first collecting space and the first impingement cooling space.
  • This supply channel preferably extends over the entire span of the airfoil. It can, more preferably, taper from its foot-side end to the head-side end, so that, provided that the coolant is fed into the supply channel at the foot-side end, it has a larger flow cross-section at the foot-side end than at its head end. This takes into account the fact that, due to the presence of impingement cooling openings in the impingement cooling wall, the amount of coolant present in the supply channel decreases with increasing distance from the end on the foot side. The conical shape of the supply channel therefore leads to an equalization of the flow rate of the coolant along the spreading direction.
  • At least one side wall of the blade is preferably on at least one additional baffle cooling wall is provided on both side walls.
  • the first impingement cooling (the leading edge of the blade) is followed in series by the suction-side impingement cooling and the pressure-side impingement cooling, although the two other impingement cooling systems on both sides of the profile center line are connected in parallel.
  • one of the two further impingement cooling spaces is arranged on the suction side and the other of the two further impingement cooling spaces is arranged on the pressure side and a separate collecting space is connected upstream of each of these two impingement cooling spaces.
  • These can preferably be provided by providing a first separating rib.
  • the coolant pressures required in the relevant collecting spaces can be set in accordance with the local thermal load on the suction-side and pressure-side side walls so that coolants are used efficiently and locally.
  • a further cavity is provided between two collecting spaces arranged on both sides of the profile center line.
  • This further cavity is preferably separated from the collecting spaces by two second separating ribs.
  • Said cavity can be used on the one hand to reduce the size of the collecting spaces to a desired level if a certain flow velocity is to be achieved in the collecting spaces.
  • the further cavity can also be used to guide a further coolant from a head end to a foot end of the airfoil if this coolant is only to be passed through the airfoil without absorbing thermal energy, if possible.
  • Such blades can be produced in particular by means of an additive process.
  • An additive method is understood in particular to be what is known as SLM technology, which is known as “selective laser melting”.
  • SLM technology which is known as “selective laser melting”.
  • This technology also known as 3D printing technology, enables metal components to produce cavities and passage openings that are comparatively small and with exact dimensions compared to turbine blades manufactured using conventional castings.
  • a turbine blade 10 relating to the invention is shown in perspective in FIG.
  • the turbine blade 10 is according to FIG Figure 1 designed as a rotor blade.
  • the invention can also be used in a guide vane, not shown, of a guide vane.
  • the turbine blade 10 comprises a blade root 12, which is shaped like a fir tree in cross-section, and a platform 14 arranged thereon.
  • the platform 14 is followed by a blade 16 which is aerodynamically curved.
  • the airfoil 16 comprises a suction side wall 22 and a pressure side wall 24 which, in relation to a hot gas flowing around the airfoil 16, extend from a leading edge 18 to a trailing edge 20.
  • a plurality of openings 28 for blowing out coolant are provided along the rear edge 20 and are separated from one another by webs 30 arranged in between.
  • the blade 16 extends along a spanwise direction from a foot-side end 26 to a head-side end 27. When the turbine blade 10 shown is used in a gas turbine with an axial flow, the spanwise direction coincides with the radial direction of the gas turbine.
  • Figure 2 shows a sectional view through the airfoil 16 according to the section line II-II as a first embodiment of an airfoil 16 according to the invention
  • Figure 3 shows a second embodiment for this.
  • the blade 16 and its pressure-side side wall 24 and suction-side side wall 22 extend - as already explained - from the leading edge 18 starting along a profile center line 32 to the rear edge.
  • impingement cooling wall 34 In the interior of the blade 16, a first perforated, ie provided with impingement cooling openings 42, impingement cooling wall 34 is arranged at a distance from the inner surface of the leading edge 18, so that a first impingement cooling space 36 is formed between them.
  • a supply channel 38 is provided on the side of the first impingement cooling wall 34 opposite the first impingement cooling space 36. This is separated from the remaining cavity of the airfoil 16 by a first rib 40.
  • the first rib 40 extends according to the cross-sectional plane from a rib end 37 on the suction side to a rib end 37 on the pressure side and has outlet openings 39 close to the side wall for the first impingement cooling space 36.
  • impingement cooling openings 42 lying in the sectional plane are provided in the first impingement cooling wall 34 for surface cooling of the front edge 18 and the directly adjoining suction-side and pressure-side areas of the side walls 22, 24.
  • first rib 40 is followed by a first collecting space 44 which is separated from a second collecting space 48 by a second rib 46.
  • the latter is also delimited by a third rib 50, so that a third collecting space 52 adjoins it further in the direction of the rear edge.
  • the first collecting space 44 is delimited both on the suction side and on the pressure side by two further impact cooling walls 54.
  • Impingement cooling openings 42 are also arranged in these, so that first further impingement cooling spaces 56 are provided with which corresponding sections of the suction-side and pressure-side side walls 22 and 24 can be impingement-cooled.
  • Second further impingement cooling spaces 59 are separated from the second collecting space 48 by impingement cooling walls 55.
  • all further impingement cooling spaces 56, 59 are laterally delimited by projections 57 extending inward from the side walls 22, 24.
  • the second and third ribs 46, 50 merge at their rib ends 37 into the suction-side and pressure-side side wall 22, 24 and have outlet openings 39 there near the side wall.
  • impingement cooling openings 42 outlet openings 39 are present, but that more of them are distributed along the span in the corresponding walls at the corresponding position, preferably lying in a row.
  • a coolant is fed to the supply channel 38 through an opening (not shown) in the turbine blade 10 during operation. There it is distributed over the span of the blade and flows through the individual impingement cooling openings 42 of the first impingement cooling wall 34, forming air jets. The air jets impinge in a known manner on the inner surface of the leading edge and cool it as intended. The coolant then flows through the outlet openings 39 of the first rib 40, after which it strikes the projections 57 with impinging cooling and is deflected by them into the first collecting space 44. From there it flows through the first and second further impact cooling walls 54, 55 for cooling the associated side wall sections. It passes from the first and second impingement cooling spaces 56, 59 through the outlet openings of the ribs 46, 50 into the subsequent collecting spaces 48, 52.
  • the coolant After the coolant has flowed through the cascaded impingement cooling arrangement described above, it reaches the collecting space 52. From there, the coolant can be fed in a known manner and further sections of the airfoil 16 are used for cooling. It is conceivable that on the one hand it is diverted into a kind of meander cooling and then blown out through the rear edge openings 28. It is also possible that the coolant from the interior of the airfoil 16 through film cooling openings (64, Fig. 3 ) is directed to the outside. The combination of both variants can also make technical sense.
  • Figure 3 shows an alternative embodiment of the turbine blade 10 according to the invention as a second embodiment.
  • the identical features are provided with the same reference numerals, so that only the differences from the first exemplary embodiment will be discussed below.
  • separating ribs 58, 60 are provided in the interior of the airfoil 16.
  • a first separating rib 58 extends between the rib 40 and the further rib 46 along the profile center line 32.
  • the separating rib 58 divides the collecting space 44 into two collecting spaces 44a and 44b, of which the former is provided on the suction side and the latter on the pressure side.
  • Two second separating ribs 60 extend along and thus quasi parallel to the profile center line 32 between the rib 46 and the rib 50, but one of them is arranged on the suction side and one on the pressure side.
  • a further cavity 62 can be provided.
  • the further cavity 62 can be used for different purposes. For example, it is suitable for conveying part of the coolant from the foot-side end 26 of the airfoil 16 to a head-side end 27 of the airfoil 16 without this coming into contact with the comparatively hot side walls 22, 24. In this way, comparatively cool cooling air can be provided at the head-side end 27 of the airfoil, which is particularly advantageous in the case of guide vanes. It is also conceivable that the cavity 62 is hermetically sealed in order to guide the cooling air guided in the sub-collection spaces 48a, 48b closer to the impingement cooling walls 54 and the impingement cooling openings 42 arranged therein.
  • the invention thus relates to an airfoil 16 for a turbine blade 10, comprising a suction-side side wall 22 and a pressure-side side wall 24, which extend along a profile center line 32 from a common leading edge 18 to a common trailing edge 20 and in a spreading direction from a base end 26 a head end 27 extending at least partially enclosing a cavity, with a first perforated baffle cooling wall 34 provided with openings inside for impingement cooling of the leading edge 18 and at least one further perforated impingement cooling wall 54 for impingement cooling of a section of the suction-side and / or pressure-side vane wall 22, 24 is provided.
  • the impingement cooling openings 42 of the first impingement cooling wall 34 and the at least one second impingement cooling wall 54 are fluidically connected in series.

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

Claims (8)

  1. Corps (16) d'aube d'une aube (10) de turbine, comprenant un extrados (22) et un intrados (24) qui, en s'étendant, le long d'une ligne médiane de profil, d'un bord (18) avant commun à un bord (20) arrière commun et, dans une direction d'envergure, d'une extrémité (26) du côté de l'emplanture à une extrémité (27) du côté de la tête entoure, au moins en partie, une cavité, dans lequel il est prévu le long de l'envergure, à l'intérieur, une première paroi (34) de refroidissement par rebondissement, pourvue de premières ouïes (42) de refroidissement par rebondissement, pour le refroidissement par rebondissement du bord (18) avant et au moins une autre paroi (54, 55) de refroidissement par rebondissement, pourvue également d'ouïes (42) de refroidissement par rebondissement, pour le refroidissement par rebondissement d'une partie de l'extrados et/ou de l'intrados (22, 24), dans lequel les ouïes (42) de refroidissement par rebondissement de la première paroi (34) de refroidissement par rebondissement et les ouïes de refroidissement par rebondissement d'au moins une deuxième paroi (54, 55) de refroidissement par rebondissement sont montées en série en technique d'écoulement et il est prévu, entre la paroi (34, 54, 55) de refroidissement par rebondissement concernée et la face intérieure de l'extrados (22) ou intrados (24) associé un espace (36, 56, 59) de refroidissement par rebondissement et dans lequel il est prévu, en aval de l'espace (36, 56, 59) de refroidissement par rebondissement concerné, respectivement un espace (44, 48, 52) de collecte, qui est voisin immédiatement en amont de l'autre paroi (34, 54, 55) de refroidissement par rebondissement en aval,
    caractérisé en ce que l'espace (44, 48) de collecte est délimité en partie par une saillie (57), qui est refroidie par rebondissement par le fait que des ouïes (39) de sortie, proches de l'extrados ou de l'intrados, sont prévues dans une nervure (40, 46), qui s'étend, suivant un plan en section transversale, d'une extrémité (37) de nervure d'extrados à une extrémité (37) de nervure d'intrados.
  2. Corps (16) d'aube suivant la revendication 1,
    dans lequel un canal (38) d'alimentation est prévu entre le premier espace (44) de collecte et le premier espace (36) de refroidissement par rebondissement.
  3. Corps (16) d'aube suivant la revendication 1,
    dans lequel, sur au moins l'un de l'extrados (22) et de l'intrados (24) du corps (16) d'aube, est prévu ou sont prévus de préférence sur les deux (22, 24) respectivement, au moins une autre paroi de refroidissement par rebondissement.
  4. Corps (16) d'aube suivant la revendication 3,
    dans lequel l'un des deux autres espaces de refroidissement par rebondissement est disposé du côté de l'intrados et l'autre des deux autres espaces de refroidissement par rebondissement est disposé du côté de l'intrados et un espace (44a, 44b, 48a, 48b) de collecte distinct est monté en amont de chacun d'entre eux.
  5. Corps (16) d'aube suivant la revendication 4,
    dans lequel un autre espace (62) vide est prévu entre deux espaces de collecte disposés des deux côtés de la ligne médiane du profil.
  6. Corps (16) d'aube suivant l'une des revendications précédentes,
    qui est monolithique.
  7. Corps (16) d'aube suivant la revendication 6,
    qui est fabriqué selon un procédé additif.
  8. Aube (10) de turbine ayant un corps (16) d'aube suivant l'une des revendications précédentes.
EP18779293.2A 2017-09-25 2018-09-19 Aubage d'aube de turbine Active EP3658751B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102017216926 2017-09-25
PCT/EP2018/075288 WO2019057743A1 (fr) 2017-09-25 2018-09-19 Aubage d'aube de turbine

Publications (2)

Publication Number Publication Date
EP3658751A1 EP3658751A1 (fr) 2020-06-03
EP3658751B1 true EP3658751B1 (fr) 2021-07-07

Family

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EP3564484A1 (fr) * 2018-05-04 2019-11-06 Siemens Aktiengesellschaft Paroi de composant d'un composant à gaz chaud
US11512597B2 (en) * 2018-11-09 2022-11-29 Raytheon Technologies Corporation Airfoil with cavity lobe adjacent cooling passage network
US11286793B2 (en) * 2019-08-20 2022-03-29 Raytheon Technologies Corporation Airfoil with ribs having connector arms and apertures defining a cooling circuit
CN111927564A (zh) * 2020-07-31 2020-11-13 中国航发贵阳发动机设计研究所 一种采用高效冷却结构的涡轮导向器叶片
US11767766B1 (en) * 2022-07-29 2023-09-26 General Electric Company Turbomachine airfoil having impingement cooling passages

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US6036441A (en) * 1998-11-16 2000-03-14 General Electric Company Series impingement cooled airfoil
JP2002242607A (ja) 2001-02-20 2002-08-28 Mitsubishi Heavy Ind Ltd ガスタービン冷却翼
US7097426B2 (en) * 2004-04-08 2006-08-29 General Electric Company Cascade impingement cooled airfoil
US7497655B1 (en) 2006-08-21 2009-03-03 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall impingement and vortex cooling
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JP5675081B2 (ja) 2009-11-25 2015-02-25 三菱重工業株式会社 翼体及びこの翼体を備えたガスタービン
US10024171B2 (en) * 2015-12-09 2018-07-17 General Electric Company Article and method of cooling an article

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WO2019057743A1 (fr) 2019-03-28
US20200277860A1 (en) 2020-09-03
US11203937B2 (en) 2021-12-21

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