EP3231995A1 - Aube de turbine comprenant des pales de turbine et une enveloppe de pale - Google Patents

Aube de turbine comprenant des pales de turbine et une enveloppe de pale Download PDF

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Publication number
EP3231995A1
EP3231995A1 EP16164811.8A EP16164811A EP3231995A1 EP 3231995 A1 EP3231995 A1 EP 3231995A1 EP 16164811 A EP16164811 A EP 16164811A EP 3231995 A1 EP3231995 A1 EP 3231995A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
cooling fluid
blade
core
turbine
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Withdrawn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
EP16164811.8A
Other languages
German (de)
English (en)
Inventor
Björn Buchholz
Waldemar Heckel
Daniela Koch
Markus Lempke
Thorsten Mattheis
Michael Ott
Marcel SCHLÖSSER
Stefan Völker
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 EP16164811.8A priority Critical patent/EP3231995A1/fr
Publication of EP3231995A1 publication Critical patent/EP3231995A1/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Withdrawn legal-status Critical Current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
    • F01D5/147—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
    • F01D5/18—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
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • 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

Definitions

  • the invention relates to a turbine blade for a turbomachine, in particular a gas turbine, with a blade root, on which a blade platform is formed, and an airfoil, which protrudes opposite to the blade root of the blade platform, wherein provided in the interior of the airfoil, a cavity for a cooling fluid is. Furthermore, the present invention relates to a turbomachine with such turbine blades.
  • Such turbine blades are known in the prior art in different configurations and are used in turbomachines to convert the thermal and kinetic energy of a working fluid, in particular a hot gas into rotational energy.
  • a turbine blade comprises a blade root on which a blade platform is formed.
  • the turbine blade comprises an airfoil, which protrudes from the blade platform opposite to the blade root.
  • turbomachines such as gas turbines, include a housing in which a flow passage extends in an axial direction.
  • a plurality of turbine stages are arranged one behind the other in the axial direction and spaced from each other.
  • Each turbine stage includes a plurality of turbine blades that form a stator vane connected to the housing and a rotor ring connected to a rotor centrally supported and passing through the housing in the axial direction.
  • the flow channel of the turbomachine flows through a hot gas.
  • the expanding hot gas flowing through the flow channel is then deflected by the guide vanes in such a way that it optimally flows against the rotor blades arranged behind it.
  • the torque generated thereby puts the runner in rotation.
  • This rotational energy can then be converted into electrical energy, for example by means of a generator.
  • an object is to provide turbine blades, which have sufficient mechanical stability for the operation of the gas turbine even at very high temperatures of the hot gas.
  • turbine blades are provided with elaborate coating systems.
  • a further increase in the allowable inlet temperature of the hot gas during operation of the gas turbine can be achieved by cooling the turbine blades.
  • cavities for a cooling fluid are provided in their interior. Common cooling methods are, for example, the impingement cooling, in which the cooling fluid is guided so that it impinges on the wall of the blade from the inside, or the film cooling, in which the cooling fluid forms a cooling film on the outside of the blade.
  • the airfoil comprises an airfoil core and a blade airfoil mounted on the airfoil core, which surrounds the airfoil core in a form-fitting manner.
  • the invention is based on the idea of providing an airfoil core and an airfoil shroud which together form the airfoil of the turbine blade in the intended condition.
  • the blade wrap surrounds the blade core in a form-fitting manner like a skin.
  • An advantage of this construction of the airfoil is that the airfoil core is additionally protected by the airfoil cover.
  • the blade leaf sheath can be easily detached from the blade core as part of a refurbishment and replaced by a new blade sheath, which is associated with low costs.
  • the blade airfoil comprises or consists of a metal or a metal alloy, in particular a nickel-based alloy such as, for example, CM247 or SIEMET-DS.
  • the airfoil sheath has a thickness in the range of 100 .mu.m to 500 .mu.m and preferably of 200 .mu.m. With such a small thickness, the airfoil shroud has a relatively small mass relative to the airfoil core. This is particularly favorable for a blade because the Blade cover which increases only insignificantly during operation due to the effective rotational forces occurring mechanical load on the blade and the blade carrier holding it.
  • the airfoil sheath is manufactured as a separate component by means of a generative production process, in particular by means of selective laser melting (SLM).
  • SLM selective laser melting
  • Generative manufacturing processes make it possible to produce components with complex shapes and, in addition, have a low expansion, so that the desired low thicknesses of the blade airfoil can be achieved.
  • the abovementioned nickel-base alloys are particularly suitable for use in additive manufacturing processes.
  • the airfoil sheath is preferably connected to the airfoil core, wherein the connection is produced by means of conventional joining methods, in particular by means of high-temperature brazing or by shrink-fitting the airfoil sheathing onto the airfoil core.
  • Conventional joining methods such as, for example, high-temperature soldering offer the advantage that the connection between the blade airfoil and the airfoil core is easy to produce and, if required, can be released again. Although a shrunk-on blade cover can only be released by destruction from the blade core. This type of connection can be achieved without additional material and / or impairment of the substance of the airfoil core or the airfoil cover.
  • the material of the airfoil sheath has a lower coefficient of thermal expansion than the material of the airfoil core. This choice of material ensures that the form-fitting fit of the blade airfoil is maintained during operation. In the case of a shrunk-on blade sheathing, moreover, detachment of the blade shroud from the blade core during operation is prevented.
  • the airfoil may have a thermal barrier coating (TBC).
  • TBC thermal barrier coating
  • Coating systems comprising a thermal barrier coating as well as an adhesive coating (Bond Coating, BC) have been proven and widely used.
  • the thermal barrier coating comprises a ceramic material, in particular partially stabilized zirconium or gadolinium zirconate or consists thereof. Ceramic materials have a particularly high heat resistance and are good thermal insulators.
  • the thermal barrier coating is arranged on the outside of the blade blade shell.
  • an adhesive layer is first applied externally to the airfoil which enhances the adhesion of the thermal barrier coating applied in a second step.
  • the adhesive layer may for example consist of a metal such as MCrALY.
  • the blade blade shell on its outside a rough surface or microstructure on such that the thermal barrier coating is mechanically clamped to the surface or the microstructuring.
  • the blade sheet alone allows a firm adhesion of the thermal barrier coating on its outside, which is why the application of an additional adhesive layer can be omitted.
  • the airfoil cover comprises at least two segments.
  • a multi-part construction of the airfoil sheath can expand the scope of a two-part construction of an airfoil of an airfoil core and a blade airfoil.
  • Complex shapes of the airfoil core can also be provided with a blade airfoil in this embodiment of the airfoil sheath, if a corresponding one-piece airfoil shell could not be pushed onto the airfoil core.
  • the airfoil core may be manufactured by a vacuum investment casting process or a 3D printing process. Vacuum investment casting processes are proven and widely used manufacturing processes for vanes of vanes. 3D printing methods can be used in small series or complex forms of blade blades, the production of which by means of a vacuum investment casting method is impossible or at least very time-consuming and / or error-prone.
  • the airfoil core has a plurality of cooling fluid bores.
  • the cooling fluid bores connect the cavity in the interior of the airfoil core with the outside of the airfoil core to direct cooling fluid out of the cavity to form an outboard cooling film.
  • the airfoil core on its outer side at least one cooling fluid channel, which has the shape of a groove, is open to the outside and connected via at least one cooling fluid bore with the cavity.
  • Such groove-shaped cooling fluid channels on the outside of the blade core are closed by the blade cover in the intended mounted state of the blade. In this way, the cooling fluid supplied through cooling fluid bores can be guided on the inside of the blade leaf cover.
  • the airfoil core has a plurality of cooling fluid ducts, wherein at least two cooling fluid ducts are connected to each other like a net, in particular all cooling fluid ducts.
  • This configuration of the cooling fluid channels allows circulation of the cooling fluid between the airfoil core and the airfoil sheath.
  • the airfoil cover has a plurality of cooling fluid bores. Cooling fluid bores allow cooling fluid to flow from the inside of the airfoil shroud to the outside thereof.
  • a first part of the cooling fluid bores of the aerofoil sheath is aligned with corresponding cooling fluid bores of the aerofoil core, when the aerofoil sheathing surrounds the aerofoil core in a form-fitting manner.
  • the cooling fluid passes from the cavity of the airfoil core to the outside of the airfoil, whereby a conventional film cooling is ensured.
  • the cooling fluid bores of the first part of the cooling fluid bores may have diameters in the range from 500 ⁇ m to 2500 ⁇ m. Holes with diameters in this range can readily be formed by conventional drilling techniques in the airfoil sheath.
  • a second part of the cooling fluid bores of the blade airfoil may be fluidly connected to corresponding cooling fluid passages of the airfoil core if the blade airfoil surrounds the airfoil core in a form-fitting manner.
  • cooling fluid circulating in the cooling fluid passages may flow through the cooling fluid bores of the second portion of the cooling fluid bores to the outside of the airfoil to form and / or supplement a cooling film.
  • this second part of the cooling fluid bores can be flexibly positioned within the limits of the arrangement of the cooling fluid ducts as required, without modifying the airfoil core.
  • the second part of the cooling fluid bores is covered by the thermal barrier coating. Accordingly, during operation of the turbine bucket, no cooling fluid penetrates through the cooling fluid bores of the second portion of the cooling fluid bores. However, if, as a result of operational wear, the thermal barrier coating is at least partially eroded, cooling fluid may exit the exposed second portion of the coolant fluid bores to ensure emergency cooling of the eroded area. Through this Notkühlung can extend the life of a turbine blade according to the invention.
  • At least one cooling fluid bore in particular all cooling fluid bores of the second part of the cooling fluid bores, has a diameter in the range between 100 ⁇ m and 500 ⁇ m. Coolant fluid wells of such small diameters can readily be formed by generative manufacturing techniques, while they can not be made by conventional drilling techniques.
  • cooling fluid bores of the second part of the cooling fluid bores are arranged on nodes of a grid. Such regular arrangements of cooling fluid bores cause the formation of a dense and uniform cooling film in the eroded area.
  • the present invention further provides a turbomachine, in particular a gas turbine, having a housing in which extends in an axial direction a flow channel through which a hot gas flows during operation of the turbomachine, and a plurality of turbine stages, each having a vane ring and a Blade rim, wherein the turbine stages are arranged in the flow channel in the axial direction one behind the other and spaced from each other, characterized in that the blade rings and / or Leitschaufelkränze comprise turbine blades according to the invention.
  • a turbomachine in particular a gas turbine, having a housing in which extends in an axial direction a flow channel through which a hot gas flows during operation of the turbomachine, and a plurality of turbine stages, each having a vane ring and a Blade rim, wherein the turbine stages are arranged in the flow channel in the axial direction one behind the other and spaced from each other, characterized in that the blade rings and / or Leitschaufelkränze comprise turbine blades according to the
  • the turbine blades When turbomachinery with turbine blades whose blades each have an airfoil core and the blade airfoil form-fitting surrounding blade cover, the turbine blades can be easily and inexpensively rehabilitate, if necessary, by only the Schaufelblatthüllen be replaced.
  • FIGS. 1 to 6 show a turbine blade 1 for a turbomachine, not shown, in particular gas turbine, according to an embodiment of the present invention.
  • the turbine blade 1 is designed for use as a rotor blade.
  • vanes may be readily practiced in accordance with the present invention.
  • the turbine blade 1 comprises a blade root 2, on which a blade platform 3 is formed. Furthermore, the turbine blade 1 comprises an airfoil 4 which protrudes from the blade platform 3 opposite the blade root 2. In the interior of the airfoil 4 is a cavity 5 for a Cooling fluid provided.
  • the airfoil 4 comprises an airfoil core 6 and a blade airfoil 7 which is mounted on the airfoil core 6 and which surrounds the airfoil core 6 in a form-fitting manner.
  • the airfoil core 6 has a plurality of cooling fluid bores 8 which connect the cavity 5 to the outside of the airfoil core 6. Further, on the outside of the airfoil core 6, a plurality of cooling fluid channels 9 are formed, which are connected to one another like a net.
  • the cooling fluid channels 9 have the shape of grooves, are open to the outside and connected to the cavity 5 via a plurality of cooling fluid bores 8.
  • the airfoil core 6 is manufactured by means of a vacuum investment casting process, but can also be produced in a 3D printing process.
  • the blade blade cover 7 is manufactured as a separate component by means of a generative production method, in particular by means of selective laser melting (SLM). It is made of a nickel base alloy such as CM247 or SIEMET-DS, but may generally include or consist of a metal or metal alloy.
  • SLM selective laser melting
  • the airfoil 7 has a thickness of about 200 microns. It is connected to the airfoil core 6 by means of high-temperature brazing, but may also be fastened to the airfoil core 6 by means of other conventional joining methods. Alternatively, a shrinkage of the blade airfoil 7 on the airfoil core 6 comes into question.
  • the material of the blade airfoil 7 has a lower coefficient of thermal expansion than the material of the airfoil core 6. In this way, the positive connection between the airfoil 7 and the airfoil core 6 is maintained even at a high temperature.
  • the airfoil 4 has a thermal barrier coating (TBC), not shown, which may include, for example, partially stabilized zirconium or gadolinium zirconate, but may also comprise or consist of another ceramic material.
  • TBC thermal barrier coating
  • the blade airfoil 7 has a rough surface on its outside or microstructuring on.
  • the heat-insulating layer arranged on the outside of the airfoil sheath 7 is mechanically clamped to the surface or to the microstructuring.
  • an adhesive layer for example of the metal MCrAlY, on the outside of the blade leaf cover 7, with which the thermal barrier layer mechanically clamps.
  • the airfoil 7 is integrally formed, as in the FIG. 3 is shown. If, due to the complex shape of the airfoil 4, the airfoil sheath 7 could not be pushed onto the airfoil core 6, the airfoil 7 could alternatively comprise two or more segments.
  • the airfoil sheath 7 further has a plurality of cooling fluid bores 10.
  • a first part of the cooling fluid bores 10 of the blade airfoil 7 is aligned with corresponding cooling fluid bores 8 of the airfoil core 6 when the airfoil 7 surrounds the airfoil core 6 in a form-fitting manner.
  • the cooling fluid bores 10 of the first part of the cooling fluid bores 10 have diameters in the range of 500 ⁇ m to 2500 ⁇ m.
  • a second part of the cooling fluid bores 10 of the blade airfoil 7 is fluidly connected to corresponding cooling fluid ducts 9 of the airfoil core 6 when the airfoil cover 7 surrounds the airfoil core 6 in a form-fitting manner.
  • the cooling fluid bores 10 of the second part of the cooling fluid bores 10 have a diameter in the range between 100 ⁇ m and 500 ⁇ m.
  • a grid-like arrangement of cooling fluid bores 10 of the second part of the cooling fluid bores 10 is provided, which is positioned in a front and platform region of the suction side of the airfoil cover 7, as in FIGS FIGS. 3 and 4 is shown. This area is particularly affected by erosion.
  • the second Part of the cooling fluid bores 10 is covered by the thermal barrier coating.
  • FIG. 5 The Figures 5 and 6 show two states of the turbine blade 1 during assembly of the blade airfoil 7 on the airfoil core 6.
  • the airfoil 7 is pushed from above onto the airfoil core 6 ( FIG. 5 ) until it surrounds the blade core 6 in a form-fitting manner ( FIG. 6 ).
  • FIG. 7 schematically shows a portion of a turbomachine according to an embodiment of the present invention.
  • the turbomachine comprises a housing 11 in which a flow channel 12 extends in an axial direction A.
  • the turbomachine comprises a plurality of turbine stages 13, each comprising a vane ring 14 and a vane ring 15, wherein the turbine stages 13 are arranged in the flow channel 12 in the axial direction A one behind the other and spaced from each other.
  • the rotor blade rings 15 are each formed from a plurality of turbine blades 1 according to the invention.
  • the flow channel 12 is traversed by a hot gas.
  • the turbine blades 1 are each traversed by a cooling fluid.
  • the cooling fluid flows from the cavity 5 through the cooling fluid bores 8 of the airfoil core and the aligned cooling fluid bores 10 of the airfoil 7 on the outside of the airfoil 4.
  • the effluent cooling fluid forms a cooling film, which is the outside of the airfoil 4 before the flowing hot gas protects and reduces the thermal load of the airfoil 4.
  • An advantage of the turbine blade according to the invention is that the airfoil 7 can be easily replaced as needed in the context of a refurbishment without having to replace the airfoil 6 core. Another advantage is the fact that an effective emergency cooling can be realized by means of the second part of film cooling holes 10 in the blade casing 7, whose effect occurs precisely when the heat-insulating layer is at least partially removed from the outside of the airfoil 4 due to erosion.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP16164811.8A 2016-04-12 2016-04-12 Aube de turbine comprenant des pales de turbine et une enveloppe de pale Withdrawn EP3231995A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP16164811.8A EP3231995A1 (fr) 2016-04-12 2016-04-12 Aube de turbine comprenant des pales de turbine et une enveloppe de pale

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP16164811.8A EP3231995A1 (fr) 2016-04-12 2016-04-12 Aube de turbine comprenant des pales de turbine et une enveloppe de pale

Publications (1)

Publication Number Publication Date
EP3231995A1 true EP3231995A1 (fr) 2017-10-18

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EP16164811.8A Withdrawn EP3231995A1 (fr) 2016-04-12 2016-04-12 Aube de turbine comprenant des pales de turbine et une enveloppe de pale

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117823234A (zh) * 2024-03-05 2024-04-05 西北工业大学 一种陶瓷纤维层叠的双空腔气冷涡轮工作叶片结构

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3810711A (en) * 1972-09-22 1974-05-14 Gen Motors Corp Cooled turbine blade and its manufacture
US20060120869A1 (en) * 2003-03-12 2006-06-08 Wilson Jack W Cooled turbine spar shell blade construction
US7625180B1 (en) * 2006-11-16 2009-12-01 Florida Turbine Technologies, Inc. Turbine blade with near-wall multi-metering and diffusion cooling circuit
US8162617B1 (en) * 2008-01-30 2012-04-24 Florida Turbine Technologies, Inc. Turbine blade with spar and shell
US20150322800A1 (en) * 2014-05-12 2015-11-12 Honeywell International Inc. Gas path components of gas turbine engines and methods for cooling the same using porous medium cooling systems
WO2016058900A1 (fr) * 2014-10-14 2016-04-21 Siemens Aktiengesellschaft Aube de turbine munie d'un module interne et procédé de fabrication d'une aube de turbine

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3810711A (en) * 1972-09-22 1974-05-14 Gen Motors Corp Cooled turbine blade and its manufacture
US20060120869A1 (en) * 2003-03-12 2006-06-08 Wilson Jack W Cooled turbine spar shell blade construction
US7625180B1 (en) * 2006-11-16 2009-12-01 Florida Turbine Technologies, Inc. Turbine blade with near-wall multi-metering and diffusion cooling circuit
US8162617B1 (en) * 2008-01-30 2012-04-24 Florida Turbine Technologies, Inc. Turbine blade with spar and shell
US20150322800A1 (en) * 2014-05-12 2015-11-12 Honeywell International Inc. Gas path components of gas turbine engines and methods for cooling the same using porous medium cooling systems
WO2016058900A1 (fr) * 2014-10-14 2016-04-21 Siemens Aktiengesellschaft Aube de turbine munie d'un module interne et procédé de fabrication d'une aube de turbine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117823234A (zh) * 2024-03-05 2024-04-05 西北工业大学 一种陶瓷纤维层叠的双空腔气冷涡轮工作叶片结构
CN117823234B (zh) * 2024-03-05 2024-05-28 西北工业大学 一种陶瓷纤维层叠的双空腔气冷涡轮工作叶片结构

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