EP3521562A1 - Segment de carénage d'extrémité destiné à l'agencement sur une aube d'une turbomachine et aube - Google Patents

Segment de carénage d'extrémité destiné à l'agencement sur une aube d'une turbomachine et aube Download PDF

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Publication number
EP3521562A1
EP3521562A1 EP19152152.5A EP19152152A EP3521562A1 EP 3521562 A1 EP3521562 A1 EP 3521562A1 EP 19152152 A EP19152152 A EP 19152152A EP 3521562 A1 EP3521562 A1 EP 3521562A1
Authority
EP
European Patent Office
Prior art keywords
rib
shroud segment
blade
ribs
sealing tip
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.)
Granted
Application number
EP19152152.5A
Other languages
German (de)
English (en)
Other versions
EP3521562B1 (fr
Inventor
Martin Pernleitner
Klaus Wittig
Lutz Friedrich
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines 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 MTU Aero Engines AG filed Critical MTU Aero Engines AG
Publication of EP3521562A1 publication Critical patent/EP3521562A1/fr
Application granted granted Critical
Publication of EP3521562B1 publication Critical patent/EP3521562B1/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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
    • F01D5/20—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
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F05D2230/10—Manufacture by removing 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2240/307—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 tip 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
    • F05D2250/70—Shape
    • F05D2250/75—Shape given by its similarity to a letter, e.g. T-shaped
    • 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
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction

Definitions

  • the present invention relates to a shroud segment for arrangement on a blade of a turbomachine according to the preamble of claim 1 and a blade according to claim 13.
  • Rotors of compressor stages and / or turbine stages of turbomachines can be exposed to high forces, in particular centrifugal forces at high speeds. These centrifugal forces can lead to deformations or material damage to the rotors.
  • outer shrouds of high-speed turbine blades can be structurally reinforced. A possible reinforcement is an improved rigidity of the outer shroud by constructive measures.
  • An object of the present invention is to propose a shroud segment having a high rigidity. Another object of the present invention is to propose a blade with a shroud segment having high rigidity.
  • the object of the invention is achieved by a shroud segment with the features of claim 1. Furthermore, the object of the invention is achieved by a blade having the features of claim 13.
  • a shroud segment for arrangement on a blade of a turbomachine, in particular of a gas turbine, with a stiffening structure raised in relation to a shroud segment surface.
  • the stiffening structure has at least three interconnected ribs.
  • the ribs can be referred to as stiffening ribs or webs.
  • a first end region of the respective at least three ribs is connected to an upstream sealing tip of the shroud segment, wherein the end region relates to the longitudinal orientation or main axis of the ribs.
  • a second end region extending to the second, longitudinally opposite end of the Ribs is connected to a downstream sealing tip of the shroud segment.
  • the blade according to the invention comprises a shroud segment for arrangement on a blade of a turbomachine with a stiffening structure raised in relation to a shroud segment surface.
  • the stiffening structure has at least three interconnected ribs. A first end region of the at least three ribs is connected to an upstream sealing tip of the shroud segment. A second end region, which refers to the second, opposite end of the ribs, is connected to a downstream sealing tip of the shroud segment.
  • the angles between the direction of the axis of rotation of the blade and the longitudinal orientations of the ribs, each viewed in the flow direction of the turbomachine, are between zero degrees and eighty degrees.
  • the blade according to the invention is made in one piece, for example by means of a casting process or by means of a generative manufacturing process.
  • the blade according to the invention may be a compressor blade and / or turbine blade of a gas turbine.
  • Exemplary embodiments according to the invention may have one or more of the following features in any combination, provided that one or the concrete combination is not apparent to the person skilled in the art as obviously technically impossible.
  • a raised stiffening structure is a structure of material accumulations, such as ribs, ridges or the like, extending radially outwardly from the shroud segment surface.
  • the stiffening structure may be made of the same material or other material as the rest of the shroud, or parts of the remaining shroud.
  • the three interconnected ribs of the shroud of the invention form a Z-shaped rib structure.
  • a sealing tip can be referred to as Dichtfin.
  • a shroud or shroud segment may be disposed on a blade tip.
  • One, two or more sealing tips on these shrouds or shroud segments may, in the event of a rotation of the shroud or shroud segment, rub against an abradable region of an inlet seal in a housing section of the turbomachine.
  • a sealing gap can form between the shroud and the housing, which minimizes the flow losses due to a backflow or a leakage flow.
  • the shroud segment can reduce the flow around the radially outer blade tip and thereby increase the efficiency of the turbomachine.
  • the shroud segments of adjacent or adjacent blades of a rotor thereby form a continuous shroud.
  • the upstream sealing tip is referred to below as the front sealing tip and the downstream arranged sealing tip as a rear sealing tip.
  • a first end portion of a first rib is connected to a first end portion of a second rib.
  • the two interconnected end portions are arranged on the front sealing tip, in particular, they are materially connected to the sealing tip.
  • a first end portion of a third rib is offset in the circumferential direction, wherein the staggered arrangement refers to the connected end portions of the first and second rib. This first end portion of the third rib is arranged at the front sealing tip, in particular materially connected to the sealing tip.
  • the pitch of the staggered arrangement of the end portion of the third rib opposite to the connected end portions of the first and second ribs is more particularly at least the width (in the circumferential direction) of the front sealing tip arranged interconnected end portions of the first and second ribs.
  • the angle between the direction of the axis of rotation of the blade and the longitudinal orientation of the first rib may be between 0 ° and 45 °, the angle between the direction of the axis of rotation of the blade and the longitudinal orientation of the third rib between 0 ° and 45 ° and / / or the angle between the direction of the axis of rotation of the blade and the longitudinal orientation of the second rib between 30 ° and 80 °.
  • the skeleton line of the blade profile intersects - in the radial direction of view - the second rib, preferably all three ribs, (each) at an angle between 30 ° and 90 °, preferably between 45 ° and 90 °.
  • the skeleton line passes through the center of each circle, which is completely inscribed in the blade profile at a certain axial position as the maximum circle.
  • roots The respective arrangements of the end portions of the ribs at the sealing tips may be referred to as roots.
  • a second end portion of the first rib is disposed on the rear sealing tip, in particular materially connected to the sealing tip.
  • a second end portion of the second rib is connected to a second end portion of the third rib.
  • a first polygonal pocket is formed between the side surfaces of the first and second ribs, the rear sealing tip, and the shroud segment surface as the bottom surface. Furthermore, a second polygonal pocket is formed between the second and third ribs, the front sealing tip and the shroud segment surface.
  • a bag can be referred to as a depression, tub, shell or the like.
  • a polygonal bag is a bag with multiple sides. The pockets have more than three side surfaces.
  • the side surfaces of the first polygonal bag are essentially formed by the first rib, the second rib and the rear sealing tip.
  • the side surfaces of the second polygonal bag are essentially formed by the second rib, the third rib and the front sealing tip. Both pockets can each have multiple side surfaces.
  • further side surfaces may be formed at the respective roots and / or at the transition regions between the ribs and the shroud segment surface.
  • the three ribs are substantially straight in their longitudinal and major axes, respectively. In other embodiments, some or all of the major axes are curved, such as single or multiple curved.
  • the end region of the third rib located at the front sealing tip is arranged on the connecting surface of the shroud segment to the closest shroud segment.
  • the end region of the first rib located at the rear sealing tip is located in the middle third, relative to the length of the shroud segment in the circumferential direction.
  • the angle between the direction of the axis of rotation of the blade, that is, the axial direction a, and the longitudinal directions of the first and third rib, each viewed in the flow direction of the turbomachine between twenty degrees and seventy degrees, in particular between thirty degrees and fifty degrees , further in particular about forty-five degrees. In other embodiments, the angle is between the directions the rotational axis of the blade and the longitudinal directions of the first and the third rib zero degrees or approximately zero degrees.
  • the ribs have a substantially radially constant height above the circumference.
  • the shroud segment is manufactured in one piece by means of a casting method, by means of a material-removing method, in particular by means of milling, or by means of a generative manufacturing method.
  • Some or all embodiments according to the invention may have one, several or all of the advantages mentioned above and / or below.
  • the shroud segment according to the invention advantageously allows a high rigidity of the shroud strip or of the outer strip, in particular in the case of high-speed turbine blades.
  • a parameter in this context is the product AN 2 , where A is the annular region formed by the blades, in particular turbine blades, and more particularly by the downstream stage. N corresponds to the speed of the blade in the application.
  • large shroud overhangs can occur.
  • the surface taper can be very large in the radial direction from the inside out. So that these shroud overhangs do not bend too much due to possible high centrifugal forces, according to the invention ribs can be introduced into the shroud.
  • the increase in mass should be minimized by the stiffening structures designed as ribs.
  • Fig. 1 shows a shroud segment 100 according to the invention with a stiffening structure and connected to the shroud segment 100 airfoil 1 in a perspective view.
  • the stiffening structure has three interconnected ribs 3, 5, 7.
  • the ribs 3, 5, 7 extend in their longitudinal direction or along their main axes in each case from one, viewed in relation to a main flow direction 9 of the turbomachine, upstream sealing tip 11 to a sealing tip 13 arranged downstream Seal tip 11 referred to as the front sealing tip 11 and the downstream arranged sealing tip 13 as a rear sealing tip 13.
  • the front sealing tip 11 may further be referred to as a sealing tip leading edge and the rear sealing tip 13 as a sealing tip trailing edge.
  • the sealing tips 11, 13 may be referred to as sealing fins.
  • the ribs 3, 5, 7 are connected to a shroud segment surface 15.
  • a sealing gap can form between the shroud and the housing, which minimizes the flow losses due to a backflow or a leakage flow.
  • the shroud segment 100 can reduce the flow around the radially outer blade tip and thereby increase the efficiency of the turbomachine.
  • the shroud segments 100 of adjacent or adjacent blades of a rotor thereby form a continuous shroud.
  • the shroud segment 100 arranged at the radial end region of the blade 1 can basically be used for damping blade vibrations, in particular in the case of gas turbine blades for turbine stages arranged at the rear, that is to say downstream.
  • the shrouds may advantageously have the shroud segments 100 according to the invention.
  • the raised stiffening structures of the shroud segments 100 of the invention may further contribute to the reduction of stress concentrations of the shroud.
  • the three interconnected ribs 3, 5, 7 are hereinafter referred to the simplified description as the first rib 3, as the second rib 5 and third rib 7.
  • the angle between the direction of the axis of rotation of the blade, which is referred to as the axial direction a and the Haut beströmungscardi 9, and the longitudinal orientations of the three ribs 3, 5, 7, viewed in the axial direction a, is exemplary in this embodiment, approximately between thirty degrees and eighty degrees. This will be in Fig. 2 described in more detail.
  • a first polygonal-shaped pocket 17 is formed, which may be referred to as a trough-shaped depression.
  • the first polygonal-shaped pocket 17 is disposed between the connecting portions of the first rib 3 and the second rib 5, the first rib 3 and the rear sealing tip 13 and between the second rib 5 and the rear sealing tip 13.
  • a second polygonal pocket 19 is formed between the second rib 5, the third rib 7 and the front sealing tip 11.
  • the second polygonal-shaped pocket 19 is disposed between the connecting portions of the second rib 5 and the third rib 7, the second rib 5 and the front sealing tip 11 and between the third rib 7 and the front sealing tip 11.
  • the sealing tips 11, 13 extend in their entire extent from below the shroud segment surface 15, ie in the region of the radially outer blade edge of the blade 1, to above the shroud segment surface 15.
  • the engagement region is an optional Inlet seal arranged in the housing region of the turbomachine.
  • the blade according to the invention comprises at least one shroud segment 100 according to the invention, an airfoil 1 and a blade root (in FIG Fig. 1 not shown).
  • the blade can be produced as a one-piece cast component, by means of a material-removing method, in particular by means of milling, or by means of a generative production method.
  • Fig. 2 shows the shroud segment 100 according to the invention Fig. 1 in a plan view from radially outside.
  • the first rib 3, the second rib 5 and the third rib 7 extend in their Longitudinal orientations, ie in the first main axis 17 of the first rib 3, in the second major axis 19 of the second rib 5 and in the third major axis 21 of the third rib 7, from the front sealing tip 11 to the rear sealing tip 13.
  • the three ribs 3, 5 7 are aligned substantially straight along their major axes 21, 23, 25.
  • the ribs 3, 5, 7 are connected to the sealing tips 11, 13 and, in this exemplary embodiment, to the shroud segment surface 15.
  • the raised stiffening structure designed as ribs 3, 5, 7 is produced integrally with the shroud segment surface 15 and the sealing tips 11, 13, for example by means of a casting process or by means of a generative production process.
  • the in Fig. 2 shown distance between on the one hand the upstream end portions of the first rib 3, the second rib 5 (which is connected to the first rib 3 in this end region) and the third rib 9 and on the other hand, the front sealing tip 11, a compound of the ribs 3, 5, Indicate 7 in the area of the shroud segment surface 15.
  • the downstream end portions of the ribs 3, 5, 7 are connected in this view directly to the rear sealing tip 11, which indicates a connection in the radially outer region of the sealing tip 13. This is also directly in Fig. 1 recognizable.
  • connection surface 27 may be referred to as a contact surface.
  • the bonding surface 27 in this exemplary embodiment is a substantially Z-shaped joint surface 27.
  • a shroud having such Z-shaped bonding surfaces 27 may be referred to as a so-called Z-Shroud.
  • the already to Fig. 1 described polygonal-shaped pockets 17, 19 have a plurality of adjacent surfaces.
  • the surfaces are in particular side surfaces.
  • the pockets 17, 19 may have four, five, six or more side surfaces.
  • the side surfaces may be arranged, for example, at the junction of the first rib 3 and the second rib 5, at the juncture of the second rib 5 and the third rib 7 and at the junctures of the ribs 3, 5, 7 with the sealing tips 11, 13. More As three side surfaces, the rigidity of the shroud segment 100 according to the invention can advantageously increase.
  • the three angles W1, W2, W3 between the direction of the axis of rotation of the blade, that is the axial direction a, and the longitudinal orientations or main axes 21, 23, 25 of the ribs 3, 5, 7 are, in each case viewed in the flow direction 9, between zero and eighty degrees.
  • the angle W1 between the axial direction a and the major axis 21 is approximately thirty degrees (30 °)
  • the angle W2 between the axial direction a and the major axis 23 is approximately sixty degrees (60 °)
  • the angle W3 between Axial direction a and the major axis 25 about thirty degrees (30 °).
  • the end region of the first rib 3 located at the rear sealing tip 13 is arranged in the middle third L1 relative to the length L of the shroud segment 100 in the circumferential direction u.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP19152152.5A 2018-01-29 2019-01-16 Aube d'une turbomachine Active EP3521562B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102018201265.2A DE102018201265A1 (de) 2018-01-29 2018-01-29 Deckbandsegment zur Anordnung an einer Schaufel einer Strömungsmaschine und Schaufel

Publications (2)

Publication Number Publication Date
EP3521562A1 true EP3521562A1 (fr) 2019-08-07
EP3521562B1 EP3521562B1 (fr) 2021-03-24

Family

ID=65033540

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19152152.5A Active EP3521562B1 (fr) 2018-01-29 2019-01-16 Aube d'une turbomachine

Country Status (4)

Country Link
US (1) US10914180B2 (fr)
EP (1) EP3521562B1 (fr)
DE (1) DE102018201265A1 (fr)
ES (1) ES2866169T3 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3865660A1 (fr) * 2020-02-11 2021-08-18 MTU Aero Engines AG Procédé d'usinage d'une pale et pale pour turbomachine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6491498B1 (en) * 2001-10-04 2002-12-10 Power Systems Mfg, Llc. Turbine blade pocket shroud
DE102009030566A1 (de) * 2009-06-26 2010-12-30 Mtu Aero Engines Gmbh Deckbandsegment zur Anordnung an einer Schaufel
EP2402559A1 (fr) * 2010-07-01 2012-01-04 MTU Aero Engines AG Aube de turbine avec plateforme d'extrémité
US20150226070A1 (en) * 2014-02-13 2015-08-13 Pratt & Whitney Canada Corp. Shrouded blade for a gas turbine engine

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5971710A (en) * 1997-10-17 1999-10-26 United Technologies Corporation Turbomachinery blade or vane with a permanent machining datum
DE10328310A1 (de) 2003-06-23 2005-01-13 Alstom Technology Ltd Verfahren zum Modifizieren der Kopplungsgeometrie bei Deckbandsegmenten von Turbinenlaufschaufeln
WO2014105533A1 (fr) 2012-12-28 2014-07-03 United Technologies Corporation Aube de turbine renforcée avec coin coupé
FR3001758B1 (fr) 2013-02-01 2016-07-15 Snecma Aube de rotor de turbomachine
US9683446B2 (en) 2013-03-07 2017-06-20 Rolls-Royce Energy Systems, Inc. Gas turbine engine shrouded blade
US9631500B2 (en) 2013-10-30 2017-04-25 General Electric Company Bucket assembly for use in a turbine engine
EP3006673A1 (fr) 2014-10-07 2016-04-13 Siemens Aktiengesellschaft Procédé et dispositif permettant de mesurer l'usure d'interverrouillage d'aubes carénées
EP3056677B1 (fr) 2015-02-12 2019-09-04 MTU Aero Engines GmbH Aube et turbomachine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6491498B1 (en) * 2001-10-04 2002-12-10 Power Systems Mfg, Llc. Turbine blade pocket shroud
DE102009030566A1 (de) * 2009-06-26 2010-12-30 Mtu Aero Engines Gmbh Deckbandsegment zur Anordnung an einer Schaufel
EP2402559A1 (fr) * 2010-07-01 2012-01-04 MTU Aero Engines AG Aube de turbine avec plateforme d'extrémité
US20150226070A1 (en) * 2014-02-13 2015-08-13 Pratt & Whitney Canada Corp. Shrouded blade for a gas turbine engine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3865660A1 (fr) * 2020-02-11 2021-08-18 MTU Aero Engines AG Procédé d'usinage d'une pale et pale pour turbomachine
US11725518B2 (en) 2020-02-11 2023-08-15 MTU Aero Engines AG Method for machining a blade and a blade for a turbomachine

Also Published As

Publication number Publication date
EP3521562B1 (fr) 2021-03-24
DE102018201265A1 (de) 2019-08-01
US10914180B2 (en) 2021-02-09
ES2866169T3 (es) 2021-10-19
US20190234219A1 (en) 2019-08-01

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