EP2639410A2 - In-situ Spielsteuerung zwischen Gasturbinenlaufschaufel und Gehäuse - Google Patents

In-situ Spielsteuerung zwischen Gasturbinenlaufschaufel und Gehäuse Download PDF

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Publication number
EP2639410A2
EP2639410A2 EP13158077.1A EP13158077A EP2639410A2 EP 2639410 A2 EP2639410 A2 EP 2639410A2 EP 13158077 A EP13158077 A EP 13158077A EP 2639410 A2 EP2639410 A2 EP 2639410A2
Authority
EP
European Patent Office
Prior art keywords
rotor
casing
compressor
ceramic coating
rotor blades
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Withdrawn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
EP13158077.1A
Other languages
English (en)
French (fr)
Inventor
Charles Alexander Smith
John David Memmer
Tushar Sharadchandra Desai
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP2639410A2 publication Critical patent/EP2639410A2/de
Withdrawn legal-status Critical Current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
    • 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
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2280/00—Materials; Properties thereof
    • F05B2280/10—Inorganic materials, e.g. metals
    • F05B2280/103—Heavy metals
    • F05B2280/10306—Hafnium
    • 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
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/12—Light metals
    • F05D2300/125—Magnesium
    • 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
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/13—Refractory metals, i.e. Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W
    • F05D2300/135—Hafnium
    • 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
    • F05D2300/21—Oxide ceramics
    • F05D2300/2112—Aluminium oxides

Definitions

  • the present disclosure relates to gas turbine engines and, more particularly, to a method and system for minimizing the clearance between the stationary rotor casing of a compressor and the rotor blades without damaging the tips of the blades, particularly during startup and shutdown.
  • the improved rotor casing design and method for coating the casing interior increases the efficiency, long term performance and reliability of the rotor, as well as the efficiency of the gas turbine engine.
  • Gas turbine engines typically include a compressor, a plurality of combustors and a gas turbine section. Compressed ambient air from the compressor mixes with a hydrocarbon fuel fed to the combustor. The fuel and compressed air mixture are then ignited, generating high temperature, expanded combustion gases. The exhaust gas is channeled through nozzles into the gas turbine in order to extract energy from the expanded gases and produce power via an electrical generator. Most gas turbines include a rotor assembly and cooperating stator that receive and redirect the hot combustion gases from one or more of the gas combustors to produce rotational energy.
  • the compressor section upstream of the turbine typically includes a rotor assembly comprising one or more rows of circumferentially-spaced rotor blades surrounded by a casing with the blades positioned between axially-spaced rows of corresponding circumferentially-spaced stator vanes.
  • the rotor blades in the compressor are coupled to a rotating disk, with each blade extending from a base platform radially outward to the blade tip.
  • ambient air flows through the rotor assembly to be directed inward by the rotor blades and then radially outward through a plurality of shrouds.
  • the stator assembly includes a corresponding plurality of stator vanes that extend radially from a base platform out to the blade tips with an outer band for mounting the stator assembly within the casing.
  • the operating temperature of both the rotor and stator assemblies increases up to a maximum anticipated level as the compressor and gas turbine engine reach a normal running speed and steady state condition.
  • the higher metal temperatures tend to migrate from the rotor blade base toward the tip of each blade.
  • the increased operating temperature of the blades may cause the tips to weaken, fracture or even deteriorate at the distal ends, causing an inevitable increase in the annular space between the blade tips and casing (sometimes referred to as the an increased "sealing gap"). Any such increase in space between the blade tips and casing during normal operation translates into a reduction of both rotor and stator efficiency, which in turn decreases the overall compressor and engine efficiency.
  • the sealing gap between the rotor blade tips and casing of the compressor should remain as small as possible without adversely restricting gas flow or effecting free blade rotation during normal operating conditions.
  • an aspect of the invention provides a system for protecting the tips of rotor blades used in various types of rotary machines, particularly gas turbine engines and compressors.
  • An exemplary embodiment of a rotor assembly includes a plurality of circumferentially spaced-apart rotor blades with each blade extending radially outward from an inner wheel disk; a cooperating stator assembly comprising one or more rows of spaced-apart stator vanes extending between adjacent rows of rotor blades; a casing extending circumferentially around the rotor and stator assemblies forming a plurality of inner flow paths defined by the rotor blades and stator vanes; and an abradable ceramic coating applied to the casing interior surface at specified locations.
  • the abradable ceramic is applied in an amount that can be abraded to form a minimum annular gap between the inner circumferential surface of the casing and tips of the rotor blades.
  • An exemplary ceramic coating applied to the casing comprises a powder containing alumina (Al 2 O 3 ) applied in situ using, for example, a plasma spray technique.
  • the coating method according to the invention can be implemented on a wide variety of rotating assemblies, particularly compressors that include a rotor rotating about a central longitudinal axis and a plurality of blades mounted to a wheel disk that extend radially outward.
  • Most rotor assemblies also include an outer casing having a generally cylindrical shape and an inner circumferential surface spaced radially outwardly from the rotor and blades to define a narrow annular gap between the inner circumferential surface of the casing and end tips of the rotor blades.
  • the abradable coatings according to the invention are applied to select portions of the inner circumference of the casing in an amount sufficient to define a minimum annular gap following abrasion as defined by the inner casing circumference and tips of the rotary blades.
  • the coating on the casing will be abraded due to slight contact with the moving blade tips as the engine and compressor reach their normal operating speed. Thereafter, the moving blade tips no longer impact the casing and the clearance between the casing and moving rotor blades becomes fixed a minimum threshold amount.
  • the protective ceramic coating is applied to the rotor casing using a plasma spray technique.
  • the coating comprises an abradable ceramic capable of being abraded by the rotor blades without damaging the blades as the compressor approaches its normal operating speed and the temperature of the system increases over time.
  • the ceramic coating is applied in situ, for example when the gas turbine engine is shut down for routine maintenance or before engine/compressor start-up. As the engine is started, the rotor blade tips contact the coated rotor casing only at the tips thereof and only a prescribed portion of the abradable ceramic applied to the casing is abraded as the compressor section reaches a steady state condition as the temperature of the blades and casing increases.
  • Exemplary abradable ceramic materials useful in practicing the invention include both "structured” and “non-structured” compositions including, by way of non-limiting example, ceramics applied in the form of a spray powder containing alumina (Al 2 O 3 ).
  • Other potentially useful ceramics include hafnia (Hf 2 ), ceria (CeO 2 ), magnesia (MgO), Yttria (Y 2 O 3 ), magnesium aluminate (MgO-Al 2 O 3 ) and zirconium silicate (ZrO 2 -SiO 2 ).
  • the powders are applied using plasma spray, chemical vapor deposition or a comparable thermal spray technique while the engine is shut down and idle.
  • the preferred thickness for ceramic layers applied to the compressor casing varies depending on the end use involved, including the aerodynamic design of the rotor blades and rotor assembly, the maximum anticipated operating temperatures of the blades, and the composition and maximum temperature of air being fed into and through the rotor and stator assemblies.
  • the ceramic layer is applied in a thickness of about 4-8 mils up to a maximum of about 20 mils.
  • the ceramic coating ensures that the desired minimum clearance will be maintained between the blades and casing while also serving as a protective layer against abrasion and a thermal barrier coating. The coatings tend to cool the rotor casing slightly and thus indirectly decrease thermal gradients in the rotor assembly.
  • the surfaces of the rotor blades and stator vanes should be roughened slightly using, for example, grit blasting to increase the adherence of the bond coat when applied by a plasma spray technique. It has been found that the level of roughness on the surface should be about 100 micro inches RMS (root mean square) to achieve the best results in applying the ceramic coatings to selected portions of the rotor casing. In some instances, it may also be useful to include additional granular particles consisting of a different, harder, ceramic material (such as corundum) to provide a more controlled and structurally sound ceramic coating that remains stable at higher anticipated operating temperatures once the compressor and engine reach a steady state condition.
  • abradable ceramics as described above can be applied to other components of the gas turbine engine, as well as to other forms of rotating equipment that rely on rotating components inside a rotor or stator assembly. Apart from the rotor assembly, the invention could be used to protect and minimize the clearance of stator vanes in the stator assembly.
  • the casing and rotor blade tip clearance control system can be carried out in situ, that is during a period of gas turbine engine down time (such as for routine, scheduled maintenance).
  • the new coating system results in significantly tighter rotor blade clearances, reduces the "secondary flow" around the blade tips and substantially improves turbine (and/or compressor) efficiency.
  • the tips of the blades become precisely positioned relative to the casing as the engine is started. This controlled "run-in” process slowly increases the clearance to an exact degree as the compressor and engine reach their anticipated high speed operating conditions.
  • FIG. 1 is schematic illustration of the major working components of on an exemplary gas turbine engine 10 coupled to electric generator 16, including a compressor section 12 that directly benefits from the in situ blade clearance control.
  • Gas turbine engine 10, compressor 12, turbine 14 and generator 16 are depicted in a single monolithic configuration with shaft 18.
  • Shaft 18 can be segmented into a plurality of segments, wherein each segment is coupled to an adjacent shaft segment.
  • Compressor 12 supplies compressed air to a combustor 20 where the air is mixed with fuel 22.
  • engine 10 could be a 6C type gas turbine engine commercially available from the General Electric Company in Greenville, S.C. In operation, air flows through compressor 12 and compressed air is supplied to combustor 20. Combustion gases 28 from combustor 20 propels turbine 14 which rotates shaft 18, compressor 12, and electric generator 16 about a common longitudinal axis.
  • FIG. 2 is a cross sectional view of the major components of an exemplary gas turbine engine labeled as shown, including rotor and stator assemblies for the compressor section, illustrating the relative location of the abradable ceramic coatings according to the invention as applied to selected portions of the compressor rotor casing.
  • FIG. 2 thus illustrates the general location of the rotor blades and stator vanes relative to the rim surfaces of the wheel disks and casing, all of which directly benefit from the abradable coating system described above as a result of the narrow gas flow path created between the casing and rotor blade tips following abrasion.
  • FIG. 3 is another cross-sectional illustration of a portion of a rotating machine (such as a compressor or turbine) that includes exemplary rotor and stator assemblies and ceramic coatings applied in the manner described herein.
  • FIG. 3 illustrates the orientation of adjacent blades and stator vanes relative to the coated casing assembled within the compressor section.
  • Compressor 30 includes a rotor assembly and a stator assembly positioned within casing 36 to define a general gas flow path 38.
  • the rotor assembly also defines an inner flow path boundary 40 of flow path 38, while the stator assembly defines an outer flow path boundary 42 of flow path 38.
  • Compressor 30 includes a plurality of stages, with each stage including a row of circumferentially-spaced rotor blades 50 and a row of stator vane assemblies 52.
  • rotor blades 50 are coupled to a rotor disk 54 with each rotor blade extending radially outwardly from rotor disk 54.
  • Each blade includes an airfoil that extends radially from an inner blade platform 58 to rotor blade tip 60.
  • the stator assembly includes a plurality of rows of stator vane assemblies 52 with each row of vane assemblies positioned between adjacent rows of rotor blades.
  • the compressor stages are configured to cooperate with a gas working fluid, such as ambient air, with the fluid being compressed in succeeding stages.
  • Each row of vane assemblies 52 includes a plurality of circumferentially-spaced stator vanes that each extend radially inward from stator casing 36 and includes an airfoil that extends from an outer vane platform 70 to a vane tip 72.
  • Each airfoil includes a leading edge and a trailing edge as shown.
  • FIG. 4 illustrates how the abradable ceramic coatings according to the invention can be applied to selected portions of the casing interior surface, including the rim surfaces of the wheel disks connected to adjacent stator vanes.
  • a plurality of rotor blades and stator vanes 80 are shown in cross section. Each of the two rotor blades 85 and 86, respectively, will abrade a portion of the ceramic applied to the casing as described above, namely ceramic coatings 81 and 88. Each blade is connected to corresponding wheel disks 82 and 87, respectively. If desired, a comparable uniform coating of abradable ceramic could be applied to rim surface 89 adjacent to stator vane 83.
  • the rotation of the compressor will abrade a precise amount of the ceramic proximate the peripheral edges of the blade tips, casing and rim surfaces in the manner described above.
  • the abrasion will continue and increase slightly as heat is generated and conducted into and along each of the rotor blades and stator vanes until the turbine or compressor reaches a steady state condition.
  • the abrasion, as controlled, ensures that a narrow sealing gap will eventually exist between the various moving rotor and stator components, thus ensuring that very little air leakage occurs between the blades, vanes and casing.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Coating By Spraying Or Casting (AREA)
EP13158077.1A 2012-03-12 2013-03-07 In-situ Spielsteuerung zwischen Gasturbinenlaufschaufel und Gehäuse Withdrawn EP2639410A2 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/417,345 US20130236302A1 (en) 2012-03-12 2012-03-12 In-situ gas turbine rotor blade and casing clearance control

Publications (1)

Publication Number Publication Date
EP2639410A2 true EP2639410A2 (de) 2013-09-18

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP13158077.1A Withdrawn EP2639410A2 (de) 2012-03-12 2013-03-07 In-situ Spielsteuerung zwischen Gasturbinenlaufschaufel und Gehäuse

Country Status (5)

Country Link
US (1) US20130236302A1 (de)
EP (1) EP2639410A2 (de)
JP (1) JP2013189977A (de)
CN (1) CN103307010A (de)
RU (1) RU2013110458A (de)

Cited By (1)

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DE102013212741A1 (de) * 2013-06-28 2014-12-31 Siemens Aktiengesellschaft Gasturbine und Hitzeschild für eine Gasturbine

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US20150075265A1 (en) * 2013-09-17 2015-03-19 General Electric Company Measurement device and method for evaluating turbomachine clearances
EP3052812A4 (de) * 2013-09-30 2016-10-05 United Technologies Corp Verdichterbereichsteilungen für einen getriebeturbolüfter
WO2016059348A1 (fr) * 2014-10-15 2016-04-21 Snecma Ensemble rotatif pour turbomachine comprenant une virole de rotor auto-portee
JP2020509228A (ja) * 2017-02-07 2020-03-26 エリコン メテコ アクチェンゲゼルシャフト、ヴォーレン アブレイダブル・コーティング
US10815783B2 (en) * 2018-05-24 2020-10-27 General Electric Company In situ engine component repair
GB201813079D0 (en) * 2018-08-10 2018-09-26 Rolls Royce Plc Effcient gas turbine engine
GB201813083D0 (en) * 2018-08-10 2018-09-26 Rolls Royce Plc Efficient gas turbine engine
US11434777B2 (en) 2020-12-18 2022-09-06 General Electric Company Turbomachine clearance control using magnetically responsive particles

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Also Published As

Publication number Publication date
RU2013110458A (ru) 2014-09-20
CN103307010A (zh) 2013-09-18
JP2013189977A (ja) 2013-09-26
US20130236302A1 (en) 2013-09-12

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