WO2014074370A2 - Bout de pale de turbine revêtu d'un abrasif - Google Patents

Bout de pale de turbine revêtu d'un abrasif Download PDF

Info

Publication number
WO2014074370A2
WO2014074370A2 PCT/US2013/067592 US2013067592W WO2014074370A2 WO 2014074370 A2 WO2014074370 A2 WO 2014074370A2 US 2013067592 W US2013067592 W US 2013067592W WO 2014074370 A2 WO2014074370 A2 WO 2014074370A2
Authority
WO
WIPO (PCT)
Prior art keywords
tip
blade
turbine
radially outer
abrasive particle
Prior art date
Application number
PCT/US2013/067592
Other languages
English (en)
Other versions
WO2014074370A3 (fr
Inventor
David B. Allen
Original Assignee
Siemens Energy, Inc.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens Energy, Inc. filed Critical Siemens Energy, Inc.
Priority to RU2015117053A priority Critical patent/RU2015117053A/ru
Priority to JP2015540755A priority patent/JP6067869B2/ja
Priority to CN201380057977.0A priority patent/CN104769228A/zh
Priority to EP13799137.8A priority patent/EP2917503A2/fr
Publication of WO2014074370A2 publication Critical patent/WO2014074370A2/fr
Publication of WO2014074370A3 publication Critical patent/WO2014074370A3/fr
Priority to IN3259DEN2015 priority patent/IN2015DN03259A/en

Links

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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/22—Non-oxide ceramics
    • F05D2300/226—Carbides

Definitions

  • This invention is directed generally to turbine blades, and more particularly to airfoil tips for turbine blades.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit.
  • Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures.
  • turbine blades are formed from a root portion at one end and an elongated portion forming a blade that extends outwardly from a platform coupled to the root portion at an opposite end of the turbine blade.
  • the blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge.
  • the tip of a turbine blade often has a tip feature to reduce the gap between ring segments and blades in the gas path of the turbine.
  • the tip features are often referred to as squealer tips and are frequently incorporated onto the tips of blades to help reduce pressure losses between turbine stages. These features are designed to minimize the gap between the blade tip and the ring segment.
  • Turbine blade tips are often coated with abrasive blade treatments to provide the turbine blades with startup cutting capacity.
  • a turbine airfoil abradable coating system with a squealer tip having a coating including an abrasive that is capable of withstanding the high temperatures of a hot gas path is disclosed.
  • the squealer tip may be attached to a radially outer surface of the tip and may be formed from at least one support materia! including at least one abrasive particle formed from a refractory carbide material that has a better resistance to thermal degradation compared to current blade tip abrasive materials as well as negligible chemical reaction with metal elements in the metal matrix used to attach the abrasive, in at least one embodiment, the abrasive particle may be tantalum carbide.
  • the squealer tip may also extend radially outward from the tip and may cover at least a portion of the radially outer surface of the tip.
  • the squealer tip of the turbine airfoil abradable coating system may be a component of a generally elongated blade having a leading edge, a trailing edge, a tip at a first end, and a root coupled to the blade at an end generally opposite the first end for supporting the blade and for coupling the blade to a disc.
  • a stationary ring segment may be positioned radially outward from the tip of the generally elongated blade and secured in position such that an inner surface of the stationary ring segment is positioned in close proximity to the generally elongated blade to limit combustion gases passing between the stationary ring segment and the generally elongated blade.
  • the inner surface of the stationary ring segment may be coated with a thermal barrier coating.
  • the squealer tip may be attached to a radially outer surface of the tip and formed from at least one support material including at least one abrasive particle formed at least partially from tantalum carbide, in another embodiment, the abrasive particle may be formed only from tantalum carbide. In yet another embodiment, the squealer tip may include more than one abrasive particle and may include a plurality of abrasive particles. The squealer tip may extend radially outward from the tip and may cover at least a portion of the radially outer surface of the tip.
  • the squealer tip may be attached to a radially outer surface of the tip and may be formed from at least one support material including at least one abrasive particle.
  • the squealer tip may extend radially outward from the tip and may cover at least a portion of the radially outer surface of the tip. In at least one embodiment, the squealer tip may cover all of the radially outer surface of the tip.
  • the squealer tip may be coated with a metal matrix used to attach the abrasive particles to the tip of the generally elongated blade.
  • the metal matrix may be formed at least partially from a metallic material, such as, but not limited to, any high temperature oxidation resistant material, such as cobalt, nickel, aluminum, chromium and yttrium (CoNiCrA!Y), In another embodiment, the metal matrix may be formed only of CoNiCrAIY.
  • An advantage of this invention is that combination of tantalum carbide particles having a size between 100 and 200 microns in size and CoNiCrAIY create an abrasive blade tip treatment with high survivability through 2,000 hours of service at conventional IGT Row 1 turbine blade tip temperatures, thereby improving clearance control for longer turbine engine service times.
  • Another advantage of this invention is that the turbine airfoil abradable coating system increases the overall engine efficiency due to its superior resistance to thermal degradation compared to the currently available abrasive blade tips.
  • Another advantage of this invention is that the turbine airfoil abradable coating system reduces the risk of blade tip wear due to tip rubbing against a radially outward stationary ring segment.
  • Figure 1 is a perspective view of a turbine blade with a squealer tip attached thereto.
  • Figure 2 is a detailed cross-sectional view of the abrasive particles within the metal matrix forming the squealer tip attached to a turbine airfoil and positioned in close proximity to a radially outward ring segment.
  • Figure 3 is a screen shot of photographs of diffusion couple preparation testing the turbine airfoil abradable coating system.
  • Figure 5 is a further magnified view from Figure 4 of intact TaC particles with no visible reaction in CoNiCrAIY matrix after 2,000 hours at 1 ,010 degrees Celsius.
  • Figure 6 is a scanning electron microscope backscattered image showing intact particles and no visible reaction with CoNiCrAIY after 2,000 hours.
  • Figure 7 is an energy dispersive spectroscopy dot map showing elemental concentrations in TaC particles with slight yttium and cobalt penetration but no loss of tantalum.
  • a turbine airfoil abradable coating system 8 with a squealer tip 10 having a coating 12 including an abrasive 14 is disclosed.
  • the squealer tip 10 may be attached to a radially outer surface 16 of the airfoil tip 24 and may be formed from at least one support material 18 including at least one abrasive particle 14 formed from a refractory carbide material that has a better resistance to thermal degradation compared to conventional blade tip abrasive materials as well as negligible chemical reaction with metal elements in the metal matrix used to attach the abrasive 14.
  • the abrasive particle 14 may be tantalum carbide.
  • the squealer tip 10 may also extend radially outward from the tip 24 and may cover at least a portion of the radially outer surface 18 of the tip 24.
  • the squealer tip 10 may be attached to a radially outward tip 24 of a turbine blade 20.
  • the turbine blade 20 may be formed from a generally elongated blade 26 having a leading edge 28, a trailing edge 30, a tip 24 at a first end 32, and a root 34 coupled to the blade 20 at a second end 36 generally opposite the first end 32 for supporting the blade 20 and for coupling the blade 20 to a disc.
  • At least one stationary ring segment 40 may be positioned radially outward from the tip 24 of the generally elongated blade 26 and secured in position such that an inner surface 42 of the stationary ring segment 40 is positioned in close proximity to the generally elongated blade 26 to limit combustion gases passing between the stationary ring segment 40 and the generally elongated blade 26.
  • the inner surface 42 of the stationary ring segment 40 may be coated with a fugitive material, such as, but not limited to, a thermal barrier coating 44.
  • the thermal barrier coating 44 may be any appropriate material that protects the stationary ring segment 40, such as, but not limited to, a porous ceramic coating.
  • the stationary ring segment 40 may be positioned such that a gap 38 exists between the outermost surface 46 of the turbine blade 20 and the stationary ring segment 40 to prevent the blade from contactsng the stationary ring segment 40. However, the size of the gap 38 is minimized to limit engine inefficiencies.
  • the squealer tip 10 may be attached to a radially outer surface 44 of the tip 24 and may be formed from at least one support material 18 including at least one abrasive particle 14.
  • the squealer tip 10 may extend radially outward from the tip 24 and may cover at least a portion of the radially outer surface 16 of the tip 24. In at least one embodiment, the squealer tip 10 may cover all of the radially outer surface 16 of the tip 24.
  • the squealer tip 10 may include a refractory carbide material that has superior resistance to thermal degradation compared to conventional blade tip abrasive materials.
  • the metal matrix 18 may support the refractory carbide material, which may have only negligible chemical reactions with the metal matrix 18.
  • the abrasive particle 14 may be formed at least partially from tantalum carbide. In another embodiment, the abrasive particle 14 may be formed only from tantalum carbide.
  • the metal matrix 18 may include one abrasive particle 14 or a plurality of abrasive particles 14.
  • the tantalum carbide particles 14 may be between 100 and 200 microns in size. Such size increases the efficiency of the squealer tip 10 by enhancing the durability during the first 2,000 hours of service use and thereafter when the capability of tip rub during warm startups exists within large IGT engines that are built on site and typically do not have a run-in procedure.
  • Diffusion couples were produced by mixing an TaC abrasive offered by American Elements, product number TA-C-02-GR with CoNiCrAIY powder offered by Praxair, part number Co-512-2 in a 50:50 ratio.
  • the mixed powder was placed into 0.5 milliliter alumnia crucibles for furnace exposure.
  • the furnace exposure was at 1 ,010 degree Celsius in air for 100, 500, 2,000 and 4,000 hours.
  • the cycle was 22 hours of heat and two hours of cooling to room temperature each day.
  • the crucible containing the TaC/CoNiCrAIY powder was removed from the furnace and epoxy was vacuum impregnated into the crucible to form a dense compact before sectioning and polishing.
  • the crucible and contents were mounted, polished and examined via a scanning electronic microscope, using energy dispersive spectroscopy to determine if any chemical reactions had occurred between the TaC and matrix elements.
  • Figures 3-6 show that the TaC grains are intact with almost no degradation after 2,000 hours at 1 ,010 degrees Celsius in air.
  • Figure 7 shows that there is little or no chemical reaction between the TaC and the CoNiCrAIY particles, indicating that the TaC material will be chemically stable in the CoNiCrAIY matrix used to plate the abrasive particles 14 to the turbine blade tips 24, such as via a TR!BOMET plate.
  • the squealer tip 10 with the metal matrix support material 18 and the abrasive particle 14 possess excellent cutting ability during initial (cold ) startup of a turbine engine, thereby enabling the turbine blade 20 to wear in with the abradable ceramic coating on the stationary ring segment 40.
  • the metal matrix support material 18 and the abrasive particle 14 is configured to survive for at least 2,000 hours while exposed to the hot gas path temperatures at a row 1 blade location in an IGT engine.

Landscapes

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

Abstract

La présente invention concerne un bout aminci qui présente un revêtement comprenant un abrasif. Le bout aminci peut être attaché à une surface radialement extérieure du bout d'aube et formé à partir d'au moins un matériau de support comprenant au moins une particule abrasive formée à partir d'un matériau de carbure réfractaire qui présente une meilleure résistance à la dégradation thermique que des matériaux abrasifs traditionnels de bout de pale, ainsi qu'une réaction chimique négligeable aux éléments métalliques dans la matrice métallique utilisée pour attacher l'abrasif. Dans au moins un mode de réalisation, la particule abrasive peut être du carbure de tantale. Le bout aminci peut aussi s'étendre radialement vers l'extérieur depuis le bout et couvrir au moins une partie de la surface radialement extérieure du bout.
PCT/US2013/067592 2012-11-06 2013-10-30 Bout de pale de turbine revêtu d'un abrasif WO2014074370A2 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
RU2015117053A RU2015117053A (ru) 2012-11-06 2013-10-30 Венец лопатки турбины с абразивным покрытием
JP2015540755A JP6067869B2 (ja) 2012-11-06 2013-10-30 タービンエアロフォイルのアブレイダブル皮膜システムおよび対応するタービンブレード
CN201380057977.0A CN104769228A (zh) 2012-11-06 2013-10-30 涡轮机翼片可磨耗涂层系统以及相应的涡轮机轮叶
EP13799137.8A EP2917503A2 (fr) 2012-11-06 2013-10-30 Bout de pale de turbine revêtu d'un abrasif
IN3259DEN2015 IN2015DN03259A (fr) 2012-11-06 2015-04-17

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201213669489A 2012-11-06 2012-11-06
US13/669,489 2012-11-06

Publications (2)

Publication Number Publication Date
WO2014074370A2 true WO2014074370A2 (fr) 2014-05-15
WO2014074370A3 WO2014074370A3 (fr) 2014-07-17

Family

ID=49684067

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2013/067592 WO2014074370A2 (fr) 2012-11-06 2013-10-30 Bout de pale de turbine revêtu d'un abrasif

Country Status (6)

Country Link
EP (1) EP2917503A2 (fr)
JP (1) JP6067869B2 (fr)
CN (1) CN104769228A (fr)
IN (1) IN2015DN03259A (fr)
RU (1) RU2015117053A (fr)
WO (1) WO2014074370A2 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2971533A4 (fr) * 2013-03-15 2016-11-16 United Technologies Corp Traitement de bout de pale de turbine pour des turbines à gaz industrielles
US10544699B2 (en) 2017-12-19 2020-01-28 Rolls-Royce Corporation System and method for minimizing the turbine blade to vane platform overlap gap
WO2024120707A1 (fr) 2022-12-05 2024-06-13 Siemens Energy Global GmbH & Co. KG Procédé d'application d'un recouvrement de blindage abrasif et protecteur et outil

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4854196A (en) * 1988-05-25 1989-08-08 General Electric Company Method of forming turbine blades with abradable tips
CA2048804A1 (fr) * 1990-11-01 1992-05-02 Roger J. Perkins Bouts d'aube mobile abrasifs longue duree
US5359770A (en) * 1992-09-08 1994-11-01 General Motors Corporation Method for bonding abrasive blade tips to the tip of a gas turbine blade
US6190124B1 (en) * 1997-11-26 2001-02-20 United Technologies Corporation Columnar zirconium oxide abrasive coating for a gas turbine engine seal system
US6641907B1 (en) * 1999-12-20 2003-11-04 Siemens Westinghouse Power Corporation High temperature erosion resistant coating and material containing compacted hollow geometric shapes
SG72959A1 (en) * 1998-06-18 2000-05-23 United Technologies Corp Article having durable ceramic coating with localized abradable portion
JP2002256808A (ja) * 2001-02-28 2002-09-11 Mitsubishi Heavy Ind Ltd 燃焼エンジン、ガスタービン及び研磨層
EP1715140A1 (fr) * 2005-04-21 2006-10-25 Siemens Aktiengesellschaft Aube de turbine ayant une bande couvrante et une couche de protection sur la bande couvrante
DE102009060570A1 (de) * 2009-12-23 2011-07-28 Lufthansa Technik AG, 22335 Verfahren zum Herstellen einer Rotor/Statordichtung einer Gasturbine

Non-Patent Citations (1)

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Title
None

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2971533A4 (fr) * 2013-03-15 2016-11-16 United Technologies Corp Traitement de bout de pale de turbine pour des turbines à gaz industrielles
US9926794B2 (en) 2013-03-15 2018-03-27 United Technologies Corporation Turbine blade tip treatment for industrial gas turbines
US10544699B2 (en) 2017-12-19 2020-01-28 Rolls-Royce Corporation System and method for minimizing the turbine blade to vane platform overlap gap
WO2024120707A1 (fr) 2022-12-05 2024-06-13 Siemens Energy Global GmbH & Co. KG Procédé d'application d'un recouvrement de blindage abrasif et protecteur et outil

Also Published As

Publication number Publication date
CN104769228A (zh) 2015-07-08
IN2015DN03259A (fr) 2015-10-09
WO2014074370A3 (fr) 2014-07-17
EP2917503A2 (fr) 2015-09-16
JP6067869B2 (ja) 2017-01-25
RU2015117053A (ru) 2016-12-27
JP2016500137A (ja) 2016-01-07

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