EP2584146A1 - Procédé de fabrication d'une aube de rotor d'une turbomachine et aube de rotor correspondante - Google Patents

Procédé de fabrication d'une aube de rotor d'une turbomachine et aube de rotor correspondante Download PDF

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Publication number
EP2584146A1
EP2584146A1 EP11186202.5A EP11186202A EP2584146A1 EP 2584146 A1 EP2584146 A1 EP 2584146A1 EP 11186202 A EP11186202 A EP 11186202A EP 2584146 A1 EP2584146 A1 EP 2584146A1
Authority
EP
European Patent Office
Prior art keywords
blade
hollow
support structure
hollow blade
area
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Withdrawn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
EP11186202.5A
Other languages
German (de)
English (en)
Inventor
Michael Clossen-Von Lanken Schulz
Kai Kadau
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 EP11186202.5A priority Critical patent/EP2584146A1/fr
Publication of EP2584146A1 publication Critical patent/EP2584146A1/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Withdrawn legal-status Critical Current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
    • 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
    • F01D5/26—Antivibration means not restricted to blade form or construction or to blade-to-blade connections or to the use of particular materials
    • 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
    • F05D2230/22—Manufacture essentially without removing material by sintering

Definitions

  • the invention relates to a method for producing a blade for a turbomachine and the blade produced by the method.
  • a gas turbine has a compressor, for example in an axial construction, in which air is compressed before it is introduced into a combustion chamber of the gas turbine for combustion.
  • the compressor has fixed vanes and rotating blades.
  • the vanes and blades are exposed to corresponding stresses during operation of the gas turbine that reduce the life of the vanes and blades.
  • the guide vanes and the blades undergo a stress due to vibrations and a centrifugal force load.
  • the centrifugal load increases with increasing length and mass of the blades and with increasing speed and is particularly high at the blade roots of the blades.
  • cracks can form, which can lead to the failure of the blade during operation of the gas turbine.
  • conventionally particularly strong materials e.g. Titanium, superalloys based on nickel or composites, used for the blades.
  • the object of the invention is to provide a method for producing a blade for a turbomachine and the blade produced by the method, wherein the blade is easily formed and yet has a high strength.
  • the method according to the invention for producing a moving blade for a turbomachine has the following steps Constructing the blade as a hollow blade whose wall thickness decreases radially outwards and dimensioned such that the blade withstands their gas force load and centrifugal load during operation of the turbomachine; Determining at least one region of the hollow blade that is excessively loaded by blade vibrations during operation of the turbomachine; Thickening the hollow blade inwardly at the area and / or providing a support structure in the interior of the hollow blade at the area, so that the strength of the hollow blade is locally increased at this area, whereby the hollow blade resists the blade vibrations during operation of the turbomachine.
  • the inventive method is particularly suitable for long blades, because in them the centrifugal load is particularly high.
  • the blade By making the blade as the hollow blade, its mass is reduced compared to a full blade, which advantageously reduces its centrifugal load during operation.
  • the wall thickness decreases radially outward, it is advantageous in the area of the blade root, ie in the region where the centrifugal force load is highest, advantageously the strongest.
  • Another advantage of the method is that the hollow blade can be produced with a low material requirement.
  • the hollow blade Due to the fact that the hollow blade has thickenings and / or the support structure on the regions which are excessively loaded by vibrations, it is advantageously reinforced selectively at these areas. At the same time, the hollow blade is made thin at less loaded areas, so that overall the hollow blade can be advantageously completed with a small mass. By the support structure can advantageously be increased, the strength of the hollow blade.
  • the method comprises the following step: Completing the blade by a generative manufacturing process.
  • the strength of the support structure decreases radially outward. It is preferred that the hollow blade is formed integrally with the support structure. This can be achieved if the hollow blade is manufactured in a single step.
  • the support structure preferably has at least one support web extending from the pressure side to the suction side of the hollow blade.
  • the support bar can thereby enclose a right angle with the wall of the hollow blade or it can form an acute angle with the wall.
  • the angle of the support bar may preferably be adapted to the forces acting on the wall such that the forces are effectively dissipated.
  • the support structure has at least two mutually crossing support webs.
  • the support structure may also preferably have a truss structure. Due to these two geometries, the strength of the hollow blade is particularly high.
  • the support structure preferably has a honeycomb structure. With the honeycomb structure can advantageously be made of a stable support structure with a low material requirement.
  • cooling air holes are arranged on the blade tip of the hollow blade. If the hollow blade cooled by cooling air from the inside, the cooling air can be advantageously dissipated by the cooling air holes at the blade tip. Furthermore, the cooling air exiting the cooling air holes blocks the radial gap at the blade tip of the built-in blade so that the leakage through the radial gap is reduced.
  • the hollow blade is preferably solid in the region near the hub.
  • the hollow blade is advantageously made particularly strong on the areas particularly heavily loaded by centrifugal forces.
  • the wall thickness preferably decreases linearly radially outward. In addition, it is preferred that the wall thickness decreases degressively in the region near the hub.
  • the additive manufacturing process is selective laser melting and / or selective laser sintering.
  • a powder of the material from which the hollow blade is produced is applied in layers to a vertically displaceable piston.
  • the areas from which the hollow blade is to be made are selectively irradiated with a laser.
  • selective laser melting the laser power is chosen so that the particles of the powder are completely melted.
  • selective laser sintering the surface of the particles is melted so far that the particles stick together after solidification of the molten surfaces. After finishing a layer, the flask is guided vertically downwards and another layer is applied.
  • the hollow blade is thus built up in layers. Due to the generative manufacturing process can advantageously produce any complex geometries of the hollow blade. In particular, the thickening and the support structures can be so easy to produce advantageous. It is also advantageously possible with the generative manufacturing method to manufacture the hollow blade in one piece and in one step together with the support structure.
  • the hollow blade has both areas produced by selective laser melting and by selective laser sintering.
  • the outsides of the hollow blade could be made by selective laser melting so that the outsides are solid.
  • the insides could be fabricated by selective laser sintering to form a porous structure with some gas permeability. This hollow blade would then be advantageous from the inside effectively cooled with cooling air. Furthermore, this could advantageously further reduce the mass of the hollow blade.
  • the blade according to the invention is produced by the method according to the invention.
  • FIG. 1 shows a perspective view of a trained as a hollow blade 1 blade of a turbomachine.
  • the hollow blade 1 is formed by a blade 2 and a platform 6, which is arranged on the hub side of the blade 2.
  • a plurality of blade cross-sections 22 to 26 of the airfoil 2 are shown, which are arranged at different radial positions of the airfoil 2.
  • the hollow blade 1 is formed by an outer wall 6 with an outer surface 7.
  • the geometry of the outer surface 7 is determined according to an aerodynamic profiling of the blade.
  • the airfoil 2 has a blade leading edge 4 and a blade trailing edge 5, and a pressure side 11 and a suction side 12.
  • the platform 9 forms in the installed state of the hollow blade 1 a hub contour, wherein on the platform 9 remote and radially outer longitudinal end of the blade 2, a blade tip 3 is formed.
  • first blade cross-section 22 The radially inner side, immediately adjacent to the platform 9 arranged, first blade cross-section 22 is solid.
  • the first blade cross-section 22 radially adjacently disposed, second blade cross-section 23 has two the inner cavity 8 bridging and secured to the outer wall 6 transverse webs 15, so that the inner cavity 8 of the hollow blade 1 is divided into three of the transverse webs 15.
  • the two transverse webs 15 are arranged substantially parallel to one another and the transverse web 15 adjoining the blade leading edge 4 in each case encloses a substantially right angle with the pressure side 11 and the suction side 12.
  • the fifth blade cross section 26 arranged radially on the outside, on the blade tip 3, is solidly formed with a plurality of cooling air holes 10 distributed over the blade tip 3.
  • FIG. 2 Two embodiments of a wall thickness curve 13 are shown in a diagram. Above the abscissa 21 is the radius and above the ordinate 20, the thickness of the outer wall 6 is plotted as the wall thickness curve 13.
  • the wall thickness profile 13 is formed continuously falling with a hub-side maximum. In the area of the hub, the wall thickness progression 13 decreases degressively in order subsequently to decrease linearly up to the blade tip 3.
  • the second embodiment of the wall thickness curve 13 differs from the first embodiment in that approximately in the radial middle position a local maximum 14 in the shape of a plateau is formed.
  • the local maximum 14 is provided at the point of the blade 2, which is excessively loaded by blade vibrations during operation of the turbomachine.
  • the thickness of the outer wall 6 according to the wall thickness curve 13 at the local maximum 14 is less than in the region of the hub, but it is also conceivable that the thickness of the outer wall 6 at the local maximum 14 is greater than in the region of the hub.
  • FIGS. 3 to 8 conceivable embodiments of the fourth blade cross-section 25 are shown. In principle, these embodiments can also be applied to the second and the third blade section 23, 24.
  • the embodiments according to FIGS. 3 to 5 and according to FIGS. 7 and 8 have a support structure 15, 16, 17, 19.
  • the embodiments according to FIGS. 6 and 7 on the other hand have inwardly directed thickening 18 of the outer wall 6.
  • the embodiment according to FIG. 3 has a plurality of pairs crossing support webs 16, which form the support structure.
  • the intersecting support webs 16 extend in the rear region of the airfoil 2 over about two-thirds of the chord length.
  • the support structure according to the in FIG. 3 shown embodiment of the crossbar 15, which bridges the mecanicholraum 8 of the hollow blade 2 from the pressure side 11 to the suction side 12 and is arranged at about one third of the chord length of the blade section 25.
  • a truss structure 17 extends as the support structure over the entire chord length of the airfoil 2, which is formed by V-shaped webs.
  • the further embodiment according to FIG. 6 has thickenings 18 both in the area of the blade leading edge 4 and in the region of the blade trailing edge 5.
  • the embodiment according to FIG. 7 has two thickenings 18 on the pressure side 11 and the suction side 12.
  • the transverse web 15 bridging the inner cavity 8 is provided, which is attached to non-thickened points of the outer wall 6.
  • the transverse web 15 is attached to the thickenings 18.
  • a honeycomb structure 19 is shown as the support structure which extends over the entire inner cavity 8.
  • the embodiment of the blade according to FIG. 7 is to be manufactured as follows: Constructing the blade as the hollow blade 1, the wall thickness radially outward according to the first embodiment of the wall thickness curve 13 according to FIG. 2 decreases linearly and is dimensioned such that the blade of their gas load and their centrifugal force load during operation of the turbomachine withstands. Next, regions of the hollow blade 1 that are excessively loaded by blade vibrations during operation of the turbomachine are to be determined. At these areas, the hollow blade 1 is to be designed with thickenings 18 which extend inwardly of the hollow blade 1, so that the strength of the hollow blade 1 is locally increased at these areas, whereby the hollow blade 1 withstands the blade vibrations during operation of the turbomachine.
  • the hollow blade 1 is to be constructed in its interior with a support structure which is formed by the transverse web 15, which bridges the pressure side 11 toward the suction side 12 in the front third of the chord length of the hollow blade 1.
  • the strength of the crosspiece 15 decreases radially outward.
  • the blade is made in one piece by selective laser sintering.
  • the manufacturing method for the embodiment of the blade according to FIG. 7 is in principle also applicable to all other described embodiments.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP11186202.5A 2011-10-21 2011-10-21 Procédé de fabrication d'une aube de rotor d'une turbomachine et aube de rotor correspondante Withdrawn EP2584146A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP11186202.5A EP2584146A1 (fr) 2011-10-21 2011-10-21 Procédé de fabrication d'une aube de rotor d'une turbomachine et aube de rotor correspondante

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP11186202.5A EP2584146A1 (fr) 2011-10-21 2011-10-21 Procédé de fabrication d'une aube de rotor d'une turbomachine et aube de rotor correspondante

Publications (1)

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EP2584146A1 true EP2584146A1 (fr) 2013-04-24

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EP11186202.5A Withdrawn EP2584146A1 (fr) 2011-10-21 2011-10-21 Procédé de fabrication d'une aube de rotor d'une turbomachine et aube de rotor correspondante

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Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102013216394A1 (de) * 2013-08-19 2015-03-12 MTU Aero Engines AG Verfahren zum Bearbeiten eines Gasturbinenbauteils
EP2851146A1 (fr) * 2013-09-24 2015-03-25 Siemens Aktiengesellschaft Procédé de fabrication d'une aube de turbine et une aube de turbine associée
WO2016128388A1 (fr) * 2015-02-11 2016-08-18 Ksb Aktiengesellschaft Composant guidant l'écoulement
EP3153663A1 (fr) 2015-10-08 2017-04-12 Rolls-Royce Deutschland Ltd & Co KG Aube de turbomachine, turboréacteur et procédé de fabrication d'une aube
WO2019042612A1 (fr) * 2017-09-01 2019-03-07 Siemens Aktiengesellschaft Aube directrice creuse pour turbines à vapeur présentant une structure intérieure produite par fabrication additive
EP3480432A3 (fr) * 2017-11-07 2019-06-05 United Technologies Corporation Treillis structurel modifié pour surfaces portantes
DE102018202194A1 (de) 2018-02-13 2019-08-14 MTU Aero Engines AG Rotorkomponente und Verfahren zum Herstellen derselben
CN110410285A (zh) * 2019-07-12 2019-11-05 湖南城市学院 智能防冻风力发电装置
FR3085418A1 (fr) * 2018-08-29 2020-03-06 Safran Aircraft Engines Aube de turbomachine comprenant une partie interne en nid d'abeille
US10871074B2 (en) * 2019-02-28 2020-12-22 Raytheon Technologies Corporation Blade/vane cooling passages
US11149550B2 (en) 2019-02-07 2021-10-19 Raytheon Technologies Corporation Blade neck transition
US11536144B2 (en) 2020-09-30 2022-12-27 General Electric Company Rotor blade damping structures
US11739645B2 (en) 2020-09-30 2023-08-29 General Electric Company Vibrational dampening elements

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3383093A (en) * 1966-06-23 1968-05-14 Gen Electric Hollow turbomachinery blades
US4934580A (en) * 1988-12-27 1990-06-19 Barnes Group, Inc. Method of making superplastically formed and diffusion bonded articles and the articles so made
GB2272731A (en) * 1992-11-18 1994-05-25 Snecma Hollow blade for the fan or compressor of a turbomachine
US6779979B1 (en) * 2003-04-23 2004-08-24 General Electric Company Methods and apparatus for structurally supporting airfoil tips
US20040191069A1 (en) * 2003-03-29 2004-09-30 Rolls-Royce Plc Hollow component with internal damping
EP1995411A2 (fr) * 2007-05-23 2008-11-26 Rolls-Royce plc Profil aérodynamique creux et son procédé de fabrication
GB2462087A (en) * 2008-07-22 2010-01-27 Rolls Royce Plc An aerofoil comprising a partition web with a chordwise or spanwise variation
EP2372092A2 (fr) * 2010-03-31 2011-10-05 General Electric Company Canaux de refroidissement pour machines industrielles

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3383093A (en) * 1966-06-23 1968-05-14 Gen Electric Hollow turbomachinery blades
US4934580A (en) * 1988-12-27 1990-06-19 Barnes Group, Inc. Method of making superplastically formed and diffusion bonded articles and the articles so made
GB2272731A (en) * 1992-11-18 1994-05-25 Snecma Hollow blade for the fan or compressor of a turbomachine
US20040191069A1 (en) * 2003-03-29 2004-09-30 Rolls-Royce Plc Hollow component with internal damping
US6779979B1 (en) * 2003-04-23 2004-08-24 General Electric Company Methods and apparatus for structurally supporting airfoil tips
EP1995411A2 (fr) * 2007-05-23 2008-11-26 Rolls-Royce plc Profil aérodynamique creux et son procédé de fabrication
GB2462087A (en) * 2008-07-22 2010-01-27 Rolls Royce Plc An aerofoil comprising a partition web with a chordwise or spanwise variation
EP2372092A2 (fr) * 2010-03-31 2011-10-05 General Electric Company Canaux de refroidissement pour machines industrielles

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102013216394B4 (de) * 2013-08-19 2015-07-23 MTU Aero Engines AG Verfahren zum Bearbeiten eines Gasturbinenbauteils
DE102013216394A1 (de) * 2013-08-19 2015-03-12 MTU Aero Engines AG Verfahren zum Bearbeiten eines Gasturbinenbauteils
EP2851146A1 (fr) * 2013-09-24 2015-03-25 Siemens Aktiengesellschaft Procédé de fabrication d'une aube de turbine et une aube de turbine associée
US11033966B2 (en) 2015-02-11 2021-06-15 Ksb Aktiengesellschaft Flow-conducting component
WO2016128388A1 (fr) * 2015-02-11 2016-08-18 Ksb Aktiengesellschaft Composant guidant l'écoulement
KR102420522B1 (ko) 2015-02-11 2022-07-13 케이에스비 에스이 앤드 코. 카게아아 유동 부품
CN107208653A (zh) * 2015-02-11 2017-09-26 Ksb 股份公司 起流动引导作用的构件
KR20170117120A (ko) * 2015-02-11 2017-10-20 케이에스비 악티엔게젤샤프트 유동 부품
CN107208653B (zh) * 2015-02-11 2022-02-18 Ksb 股份公司 起流动引导作用的构件
DE102015219530A1 (de) 2015-10-08 2017-04-13 Rolls-Royce Deutschland Ltd & Co Kg Schaufel für eine Strömungsmaschine, Turbofantriebwerk und ein Verfahren zur Herstellung einer Schaufel
EP3153663A1 (fr) 2015-10-08 2017-04-12 Rolls-Royce Deutschland Ltd & Co KG Aube de turbomachine, turboréacteur et procédé de fabrication d'une aube
WO2019042612A1 (fr) * 2017-09-01 2019-03-07 Siemens Aktiengesellschaft Aube directrice creuse pour turbines à vapeur présentant une structure intérieure produite par fabrication additive
US10378364B2 (en) 2017-11-07 2019-08-13 United Technologies Corporation Modified structural truss for airfoils
EP3480432A3 (fr) * 2017-11-07 2019-06-05 United Technologies Corporation Treillis structurel modifié pour surfaces portantes
DE102018202194A1 (de) 2018-02-13 2019-08-14 MTU Aero Engines AG Rotorkomponente und Verfahren zum Herstellen derselben
FR3085418A1 (fr) * 2018-08-29 2020-03-06 Safran Aircraft Engines Aube de turbomachine comprenant une partie interne en nid d'abeille
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
CN110410285A (zh) * 2019-07-12 2019-11-05 湖南城市学院 智能防冻风力发电装置
US11536144B2 (en) 2020-09-30 2022-12-27 General Electric Company Rotor blade damping structures
US11739645B2 (en) 2020-09-30 2023-08-29 General Electric Company Vibrational dampening elements

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