EP3216983A1 - Aube de turbine a gaz comprenant une arete de friction refroidie - Google Patents

Aube de turbine a gaz comprenant une arete de friction refroidie Download PDF

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Publication number
EP3216983A1
EP3216983A1 EP16159107.8A EP16159107A EP3216983A1 EP 3216983 A1 EP3216983 A1 EP 3216983A1 EP 16159107 A EP16159107 A EP 16159107A EP 3216983 A1 EP3216983 A1 EP 3216983A1
Authority
EP
European Patent Office
Prior art keywords
squealer
blade according
edge
recess
peripheral wall
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.)
Withdrawn
Application number
EP16159107.8A
Other languages
German (de)
English (en)
Inventor
Markus Gill
Christian Gindorf
Andreas Heselhaus
Robert Kunte
Marcel SCHLÖSSER
Andrew Carlson
Ross PETERSON
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 AG
Original Assignee
Siemens AG
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 AG filed Critical Siemens AG
Priority to EP16159107.8A priority Critical patent/EP3216983A1/fr
Priority to CN201790000656.0U priority patent/CN209976583U/zh
Priority to PCT/EP2017/054734 priority patent/WO2017153219A1/fr
Priority to US16/081,205 priority patent/US11136892B2/en
Priority to EP17707889.6A priority patent/EP3400373B1/fr
Publication of EP3216983A1 publication Critical patent/EP3216983A1/fr
Withdrawn legal-status Critical Current

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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
    • F01D5/187—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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to a rotor blade for a gas turbine, comprising a radially extending airfoil having an airfoil body having a peripheral wall with a pressure-side wall portion and a suction-side wall portion, a plate-shaped crown bottom connected in the region of the blade tip with the peripheral wall and a along the circumferential wall extending squealer, wherein the peripheral wall and the crown base define a cavity in the airfoil body, the squealer aligned on the outside with the peripheral wall and projects radially beyond the crown bottom and formed in the airfoil body cooling channels extending from the cavity to in the squealer extending cooling fluid outlet openings extend.
  • the gas turbine plant has a flow channel, in whose axial direction a turbine rotor is rotatably mounted.
  • This comprises a plurality of wheel disks, at the radially outer end surfaces of each of which a plurality of rotor blades in the form of a blade ring are arranged.
  • the rotor blades each have blade feet, which are inserted into one or more receiving grooves formed on the end faces of the wheel disks and fixed therein.
  • blade platforms are formed, from whose outer sides facing away from the wheel disc blades project into the flow channel.
  • the flow channel is traversed by the hot gas, the flowing hot gas impinging the blades with a force which, due to the shape of the airfoils, is converted to a torque acting on the turbine rotor which rotatably drives the turbine rotor.
  • an object is to provide blades that have sufficient mechanical strength for the operation of the gas turbine plant even at high thermal loads.
  • blades are provided with elaborate coating systems.
  • blades are cooled during operation of the gas turbine engine. For this purpose, in their interior cavities and cooling channels are formed, which are flowed through by a cooling fluid, usually air.
  • Common cooling methods include, for example, impingement cooling in which the cooling fluid is guided so as to impinge on the wall of the airfoil from the inside, or the film cooling in which the cooling fluid flows outwardly from the interior of the airfoil by cooling holes formed in the airfoil body on the outside of which to form a cooling film.
  • Common cast airfoils each include a hollow airfoil body which is closed in the region of the blade tip by a so-called crown bottom. Further, in the region of the blade tip, the blade body carries a squeal edge which is integrally molded on the outside of the airfoil body and protrudes along the outer contour of the peripheral wall of the airfoil body in the radial direction.
  • the turbine runner may creep out of its original central position, increase the length of the blades due to centrifugal force, or ovalize an originally circular flow channel.
  • These effects result from setting and / or elongation due to thermal stress from the hot gas and / or from rotational centrifugal forces or gravity.
  • the thus caused contact between the end faces of the squealer edges and the channel wall leads to a friction-related removal of material in the form of metal dust or metal chips from the squealer edges.
  • the cooling fluid outlet openings may become clogged with the removed airfoil material, thereby impairing or preventing cooling of the squealer edges.
  • the inadequate cooling of the squealer edges results in higher wear and consequently lower blade life.
  • the present invention provides a blade for a gas turbine of the type mentioned, in which in the end face of the squealer at least a recess is formed into which at least a part of the cooling channels opens such that the cooling fluid outlet openings lie completely in a bottom region of the at least one depression.
  • the invention is based on the consideration to lower the cooling fluid outlet openings relative to the radial direction relative to the end face of the squealer. This is achieved according to the invention in that at least one depression is formed in the end face of the squealer edge and at least part of the cooling outlet openings are arranged completely in a bottom region of the at least one depression. In this way, the cooling fluid outlet openings are removed from the contact area between the end face of the squealer edge and the channel wall, whereby a clogging of the cooling fluid outlet openings with removed airfoil material is reduced or prevented. As a result, the cooling performance over the service life of the gas turbine plant is essentially maintained, which is associated with a correspondingly long life of the blades.
  • the bottom region of the at least one depression is arranged between the end face of the squealer edge and the outside surface of the crown base, in particular the bottom region being formed as a flat bottom surface having a depth opposite to the end face Range of 0.5 mm to 4.5 mmm and preferably in the range of 0.5 mm to 2.5 mm.
  • a radial position of the bottom region causes the cooling fluid outlet openings to be arranged in the immediate vicinity of the free end region of the squealer edge, as a result of which effective cooling of this area of the squealer edge can be ensured.
  • the small depth of the bottom surface of the recess opposite the end face is sufficient to prevent material particles removed from the end face from clogging the cooling fluid outlet openings, which is accompanied by a constant cooling capacity.
  • the squeal edge relative to the radial direction relative to the outer surface of the crown base, has a height which is in the range of 1 mm to 10 mm, advantageously in the range of 1.5 mm to 6 mm, and preferably 3.5 mm. In squiggings with a height in this range, wells of suitable depth can be easily formed.
  • an inner surface of the squealer edge is inclined outward with respect to the radial direction, the angle of inclination being measured in a radially extending plane perpendicularly intersecting the squealer edge, being in the range of 0 ° to 45 °, and preferably less than 30 °. Due to the inclination of the inner surface of the squealer edge, the squealer edge widened from the end face in the direction of the crown bottom. This improves the stability of the squealer edge and additionally improves the heat transfer between the squealer edge and the crown bottom or the circumferential wall.
  • the at least one depression extends to form a stepped cross section up to the inside of the squealer edge, wherein in particular a step corner of the cross section, preferably the inner corner is rounded.
  • at least one recess is open towards the inside.
  • Such depressions can already be easily produced during the casting of the airfoil body or only later, for example, by milling or erosion.
  • the end face of the squealer edge has a width that is less than the thickness of the peripheral wall of the airfoil body in the region of the at least one depression.
  • the end face of Abrading edge have a width which is smaller than the width of the bottom portion of the at least one recess. In this way, only a relatively narrow outer region of the squealer forms its radially outer end region.
  • the end face of the squealer edge and the bottom portion of the at least one recess together have a width which is approximately equal to the thickness of the peripheral wall of the airfoil body in the region of the at least one recess.
  • Such rubbing edges essentially represent a continuation of the peripheral wall of the airfoil body beyond the crown bottom.
  • the depression in the end face of the squealer edge may be formed as a groove while leaving out an outside end face section and an inside face face section, wherein in particular the inside corners of the depression are rounded off.
  • the width of the outside end surface portion and the width of the inside end surface portion of the squeal edge may each be in the range of 0.5 mm to 5 mm, and preferably at least 1 mm, the ratio between the outside width and the outside inside width in the range between 0.7 mm and 1.3 mm, in particular 0.9 and 1.1 and is preferably 1.
  • the peripheral wall tapers in the direction of the crown bottom in favor of the cavity, the thickness of the peripheral wall being reduced from an initial thickness to a tapered thickness which is at least half the initial thickness, and the tapering over a radial portion of the peripheral wall takes place, the height of which is at least five times and at most ten times as large as the initial thickness.
  • the cooling channels may be formed so that they extend closer to the outside of the squeal edge, which is accompanied by an improved convective cooling of the squealer.
  • the cooling fluid outlet openings are arranged next to one another and at a distance from each other, in particular equidistantly and / or along a line.
  • Such arranged cooling fluid outlet openings are particularly suitable to cool the squeal along its circumferential extent. In principle, however, the cooling fluid outlet openings can be distributed as desired.
  • the at least one depression can be provided only in a section of the squealer edge projecting from the suction-side wall section of the surrounding wall. In this way, the cooling of the protruding from the suction-side wall portion of the peripheral wall portion of the squealer can be improved.
  • a plurality of recesses arranged side by side in the circumferential direction can be provided, into each of which a part of the cooling channels opens and in particular each have at least one feature mentioned above.
  • Several recesses lead to a corresponding grouping of the cooling channels.
  • each cooling channel extends in a straight line and / or has a circular cross-section with a diameter which is in the range of 0.25 mm to 2 mm and is preferably 0.6 mm.
  • the cooling channels can be widened in the region of the cooling fluid outlet openings, wherein the widenings in particular have the shape of a cylinder whose height is at most five times, preferably as large as the diameter of the cooling channel and / or its diameter at most three times, preferably twice as large as that Diameter of the cooling channel.
  • Such expanded cooling fluid outlet openings can act as a diffuser and expand the exiting cooling fluid flow accordingly, so that according to the principle of film cooling, a large portion of the squealer can be cooled.
  • the cooling fluid outlet openings can also be widened in a conical, semi-conical or fan-like manner.
  • the cooling channels are formed as bores.
  • rectilinear cooling channels of circular cross-section can be easily inserted into a cast airfoil body.
  • the cooling channels relative to the radial direction are inclined transversely to the inner surface of the squealer, in particular, the inclination angle of the cooling channels, which are each measured in a radially extending plane which intersects the squeal edge perpendicular, equal to or approximately equal to the inclination angle of the inner surface the rubbing edge are.
  • Cooling channels with such an inclination guide the cooling fluid emerging from the cooling fluid outlet openings from the inside to the outer end area of the squealer edge.
  • each cooling channel is inclined relative to a plane perpendicular to the radial direction in the direction of the leading edge of the blade or in the direction of the trailing edge of the blade, the angle of inclination in the direction of the trailing edge of the blade and the angle of inclination in the direction of the leading edge of the blade in each case
  • Level which intersects the plane of measurement of the angle of inclination perpendicular, be measured in the range between 30 ° and 90 ° lie and preferably 45 °.
  • Cooling channels with such an inclination in the direction of the leading edge or in the direction of the trailing edge have a greater length, whereby the convective cooling of the squealer can improve. In addition, they can favorably influence the flow direction of the exiting cooling fluid. Cooling channels of different inclination directions can penetrate or intersect without penetration.
  • a transition region between an inner surface of the squealer edge and the outer surface of the crown base is rounded off. This improves the aerodynamic properties of the blade tip.
  • the blade body is produced by casting or in a generative process, in particular by means of 3D printing.
  • Casting has been found to be a suitable manufacturing process, in particular for cooled airfoils with a cavity in their interior.
  • generative processes are suitable for the production of airfoil bodies.
  • FIGS. 1 to 3 show a blade for a gas turbine according to a first embodiment of the present invention.
  • the blade includes an airfoil 1 extending in a radial direction R with a cast airfoil body 2.
  • the airfoil body 2 has a peripheral wall 3 having a pressure-side wall portion 3a and a suction-side wall portion 3b.
  • the airfoil body 2 comprises a plate-shaped crown base 4, which is connected to the peripheral wall 3 in the region of the blade tip 5.
  • the peripheral wall 3 and the crown bottom 4 define in the airfoil body 2 a cavity 6, which is flowed through during the operation of the gas turbine by a cooling fluid.
  • the airfoil body 2 comprises a squealer edge 7.
  • the squealer edge 7 extends along the peripheral wall 3 and is aligned on the outside with this.
  • the squealer edge 7 projects radially beyond the crown base 4 and has, relative to the radial direction R relative to the outer surface 4a of the crown base, a height h which is measured perpendicular to the outer surface 4a of the crown base and is about 3 mm.
  • An inner surface 7a of the squealer edge 7 is inclined relative to the radial direction R by an inclination angle ⁇ of approximately 25 °, which is measured in a plane extending in the radial direction (R), which intersects the squealer edge 7 perpendicularly.
  • a transition region 8 between the inner surface 7a of the squealer edge 7 and the outer surface 4a of the crown base 4 is rounded.
  • a recess 9 is formed, which extends to form a stepped cross-section to the inside of the squealer 7.
  • the inner corner 10 of the stepped cross section is rounded.
  • the bottom portion 9a of the recess 9 is formed as a flat bottom surface and arranged with respect to the radial direction R between the end face 7b of the squealer edge 7 and the outer surface 4a of the crown bottom 4.
  • the outer surface 4a of the crown base 4, the bottom surface 9a of the recess 9 and end face 7b of the squealer 7 extend parallel to each other and perpendicular to the radial direction R.
  • the recess 9 relative to the end face 7b has a depth h 1 , as vertical distance between the bottom surface 9a and the end face 7b is measured and is about 1 mm. Accordingly, the vertically measured height h 2 of the bottom surface of the recess 9 over the outer surface 4 a of the crown base 4 is about 2 mm.
  • the bottom surface 9 a of the recess 9 and the outer surface 4 a of the crown base 4 can also be inclined to each other and / or to the radial direction R, wherein the depth h 1 or the height h 2 then in each case based on the inner corner 10 are to be determined.
  • the end face 7b of the squealer 7 has a width a 1 , which is smaller than the thickness d 1 of the peripheral wall 3 of the airfoil body 2 in the region of the recess 9.
  • the width a 1 of the end face 7b of Sharp edge 7 in the region of the recess 9 is less than the width b 1 of the bottom portion 9a of the recess 9.
  • the end face 7b of the squealer 7 and the bottom portion 9a of the recess 9 have a width a 1 + b 1 , which is approximately equal to the thickness d 1 of the peripheral wall 3 of the airfoil body 2 in the region of the recess 9, wherein the thickness d 1 is measured as a vertical distance between the outer surface and the inner surface of the surrounding wall 3.
  • the widths a 1 and b 1 respectively parallel to each other and to the outer surface 4 a of the crown bottom 4 are measured.
  • Other embodiments of the present invention may have 1, which differ from the selected here relative proportions of the widths of a 1 and b 1 as well as the thickness d.
  • cooling channels 11 are formed, which extend from the cavity 6 to cooling fluid outlet openings 12 which are provided in the squealer 7.
  • the cooling channels 11 open into the recess 9 in such a way that the cooling fluid outlet openings 12 are arranged completely in the bottom area 9a of the recess 9.
  • the cooling fluid outlet openings 12 in the recess 9 are equidistant and arranged alongside one another along a line.
  • Each cooling channel 11 is formed as a bore and extends in a straight line. It has a circular cross section with a diameter of about 0.6 mm.
  • Each cooling passage 11 is inclined with respect to the radial direction R across the inner surface 7a of the squealer edge 7, wherein the inclination angles ⁇ of the cooling channels 11 each measured in a plane extending in the radial direction R are the squealer edge 7 intersects perpendicularly, approximately equal to the angle of inclination ⁇ of the inner surface 7a of the squealer edge 7.
  • the FIG. 4 shows a blade for a gas turbine according to a second embodiment of the present invention.
  • the structure of this blade is basically the structure of the in the FIGS. 1 to 3 illustrated first embodiment.
  • the cooling channels are widened in the region of the cooling fluid outlet openings.
  • the expanded cooling fluid opening 12a has the shape of a cylinder, the height h 5 is equal to the diameter of the cooling channel 11 and the diameter c 5 is twice as large as the diameter of the cooling channel 11, resulting in a cross-sectional area for the cylinder, which is four times as large is like the cross-sectional area of the cooling channel 11.
  • a widened cooling flow is generated in accordance with which a large area of the squealer edge 7 can be cooled.
  • the FIG. 5 shows a blade for a gas turbine according to a third embodiment of the present invention. It basically has the same structure as the one in the FIGS. 1 to 3 illustrated blade.
  • the recess 9 is formed as a groove leaving an outside end face portion and an inside end face portion, so does not extend to the inside of the squeal edge 7, but is also bounded on the inside by the squeal 7.
  • the outside end face 7b have a width a 2
  • the inside end face 7b has a width c 2
  • the bottom area 9a of the recess 9 has a width b 2 .
  • the FIG. 6 shows a blade for a gas turbine according to a fourth embodiment of the present invention. It differs from the previously described embodiments in that the circumferential wall 3 tapers in the direction of the crown base 4 in favor of the cavity 6.
  • the thickness of the peripheral wall 3 is reduced from an initial thickness d 1 to a tapered thickness d 2 , which is approximately half the initial thickness d 1 .
  • the taper takes place via a radial section of the circumferential wall 3, whose height 1 is approximately five times as great as the initial thickness d 1 .
  • the taper is linear, ie the inside of the peripheral wall 3 is flat and inclined relative to embodiments without tapering of the peripheral wall 3 by an angle ⁇ .
  • the bank angle ⁇ of the cooling channels 11 is selected smaller such that the cooling channels 11 extend closer to the outside of the squealer edge 7, whereby the convective cooling of the squealer edge 7 is improved.
  • the transition region to the crown bottom 4 is rounded, wherein the curvature is defined by a radius of curvature r 2 , which may deviate from the radius of curvature r 1 of embodiments without tapering of the peripheral wall 3.
  • a radius of curvature r 2 is shown which is approximately twice as large as r 1 .
  • the transition region of the taper facing away from the crown bottom 4 is rounded to avoid an edge, wherein the rounding is defined by a radius of curvature r 3 .
  • FIG. 7 shows a blade for a gas turbine according to a fifth embodiment of the present invention. It has the same basic structure as the previously described embodiments and differs from the previously described embodiments in that the cooling channels are opposite to a plane perpendicular to the radial direction R in the direction of the trailing edge of the blade are inclined. In this case, the inclination angle ⁇ in the direction of the trailing edge of the blade in a plane which intersects the measurement plane of the inclination angle ⁇ perpendicular, measured and are 45 °. As a result, the cooling channels 11 have a greater length, as a result of which the convective cooling of the squealer edge 7 is improved.
  • FIG. 8 shows a blade for a gas turbine according to a sixth embodiment of the present invention. It is different from the one in FIG. 7 illustrated embodiments, characterized in that further cooling channels 11 are provided which are inclined relative to a plane perpendicular to the radial direction R in the direction of the leading edge of the blade.
  • the cooling channels 11 of different inclination directions penetrate each other. Alternatively, however, they can also intersect without penetration, in particular if the cooling fluid outlet openings 12 are arranged in two rows arranged next to one another.
  • the inclination angle ⁇ may be selected differently from the inclination angle ⁇ .
  • An advantage of the blade according to the invention is that the cooling channels 11 are not added or only slightly by material removal from the end face 7b of the squealer 7. This ensures a constant during operation of the gas turbine cooling the squeal edge 7 and thus a long life of the blade.
  • Another advantage of the blade according to the invention is shown in the ease of manufacture of the recess 9 and the cooling channels 11. Due to the small depth of the recess 9 is an effective cooling of the squealer 7 over its entire height h possible. In addition, the cooling fluid flowing out of the cooling fluid outlet openings 12 on its short path to the outside step of the squealer edge 7 during the Operation of the gas turbine hardly distracted, which is accompanied by an effective cooling of the blade tip 5.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP16159107.8A 2016-03-08 2016-03-08 Aube de turbine a gaz comprenant une arete de friction refroidie Withdrawn EP3216983A1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
EP16159107.8A EP3216983A1 (fr) 2016-03-08 2016-03-08 Aube de turbine a gaz comprenant une arete de friction refroidie
CN201790000656.0U CN209976583U (zh) 2016-03-08 2017-03-01 用于燃气轮机的具有冷却的扫掠边缘的转子叶片
PCT/EP2017/054734 WO2017153219A1 (fr) 2016-03-08 2017-03-01 Aube mobile pour turbine à gaz avec bord de frottement refroidi
US16/081,205 US11136892B2 (en) 2016-03-08 2017-03-01 Rotor blade for a gas turbine with a cooled sweep edge
EP17707889.6A EP3400373B1 (fr) 2016-03-08 2017-03-01 Aube de turbine a gaz comprenant une arete de friction refroidie

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP16159107.8A EP3216983A1 (fr) 2016-03-08 2016-03-08 Aube de turbine a gaz comprenant une arete de friction refroidie

Publications (1)

Publication Number Publication Date
EP3216983A1 true EP3216983A1 (fr) 2017-09-13

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EP16159107.8A Withdrawn EP3216983A1 (fr) 2016-03-08 2016-03-08 Aube de turbine a gaz comprenant une arete de friction refroidie
EP17707889.6A Active EP3400373B1 (fr) 2016-03-08 2017-03-01 Aube de turbine a gaz comprenant une arete de friction refroidie

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EP17707889.6A Active EP3400373B1 (fr) 2016-03-08 2017-03-01 Aube de turbine a gaz comprenant une arete de friction refroidie

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US (1) US11136892B2 (fr)
EP (2) EP3216983A1 (fr)
CN (1) CN209976583U (fr)
WO (1) WO2017153219A1 (fr)

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EP3477058A1 (fr) * 2017-10-24 2019-05-01 United Technologies Corporation Circuit de refroidissement de profil aérodynamique
US10436038B2 (en) 2015-12-07 2019-10-08 General Electric Company Turbine engine with an airfoil having a tip shelf outlet
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DE102020202891A1 (de) * 2020-03-06 2021-09-09 Siemens Aktiengesellschaft Turbinenschaufelspitze, Turbinenschaufel und Verfahren

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US11136892B2 (en) 2021-10-05
US20200386104A1 (en) 2020-12-10
EP3400373A1 (fr) 2018-11-14
CN209976583U (zh) 2020-01-21
WO2017153219A1 (fr) 2017-09-14
EP3400373B1 (fr) 2021-04-28

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