US20150159488A1 - Turbine rotor blade of a gas turbine and method for cooling a blade tip of a turbine rotor blade of a gas turbine - Google Patents

Turbine rotor blade of a gas turbine and method for cooling a blade tip of a turbine rotor blade of a gas turbine Download PDF

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Publication number
US20150159488A1
US20150159488A1 US14/559,668 US201414559668A US2015159488A1 US 20150159488 A1 US20150159488 A1 US 20150159488A1 US 201414559668 A US201414559668 A US 201414559668A US 2015159488 A1 US2015159488 A1 US 2015159488A1
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US
United States
Prior art keywords
blade tip
blade
turbine rotor
turbine
rotor blade
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.)
Abandoned
Application number
US14/559,668
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English (en)
Inventor
Knut LEHMANN
Anthony RAWLINSON
Jens Taege
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.)
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
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 Rolls Royce Deutschland Ltd and Co KG, Rolls Royce PLC filed Critical Rolls Royce Deutschland Ltd and Co KG
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG, ROLLS-ROYCE PLC reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: LEHMANN, KNUT, TAEGE, JENS, RAWLINSON, ANTHONY
Publication of US20150159488A1 publication Critical patent/US20150159488A1/en
Abandoned legal-status Critical Current

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    • 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
    • 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
    • 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
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • 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
    • 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
    • F05D2240/80—Platforms for stationary or moving blades

Definitions

  • This invention relates to a turbine rotor blade of a gas turbine with a blade tip. Furthermore, this invention relates to a method for cooling such a blade tip of a turbine rotor blade.
  • circumferential sealing edges are used. Designs are also known, where overhangs at the blade tip (winglets) are provided. The circumferential sealing edges can contribute to an improvement in the aerodynamics. The overhangs on the suction side and/or on the pressure side can reduce the leakage mass flow and also improve the aerodynamics around the blade tip.
  • film cooling air When film cooling air is used, it is necessary to consume a large quantity of “expensive” film cooling air in order to protect the blade tip from the high temperatures of the hot gas.
  • the film cooling air must be taken from the compressor of the gas turbine, thereby reducing the efficiency of the thermodynamic cyclic process inside the gas-turbine engine.
  • Cooling air that must be forcibly discharged at the blade tip in order to prevent dirt accumulations in the internal blade ducts contributes only to a limited extent to the cooling of the blade tip, since it cannot reach the thermally most highly loaded points of the blade tip.
  • the object underlying the present invention is to provide a method for cooling a blade tip of a turbine rotor blade of a gas turbine as well as a turbine rotor blade suitable for carrying out the method, which, while being simply designed guarantees effective cooling.
  • the invention thus provides for the thermally highly loaded rear area of the turbine rotor blade tip to be supplied at least partly with passive air, i.e. with relatively cold air from the hot gas flow. It is self-evident that the hot gas temperature must be below the maximum temperature sustainable by the blade material. There is an area with relatively low hot gas temperature on the front suction-side blade tip. In particular in the case of additional discharge of cooling air at the turbine casing upstream of the rotor blades, the hot gas temperature in this area can be considerably below the highest permissible metal temperature.
  • the relatively “cold” hot gas flow is thus used for cooling the blade tip by being routed to the thermally most highly loaded pressure-side rear area of the blade tip.
  • a cover forming a flow duct is attached at the blade tip. This creates a duct or a cavity on the blade tip, through which the cooling air can be passed.
  • This duct or cavity, forming the means in accordance with the invention for duct-type guidance of cooling air preferably has an inlet opening on the suction-side blade leading edge. It is however also possible to introduce cooling air from the blade interior. In the area of the blade trailing edge too, an opening is provided through which the cooling air flows out. The pressure difference over the blade row ensures here a flow of relatively cold hot gas through the duct or cavity.
  • the means provided in accordance with the invention can extend over the entire length or only over part of the length of the blade tip.
  • the cooling air can exit the duct or cavity in accordance with the invention at the blade tip, depending on the required counter-pressure level, in the area of the pressure-side or the suction-side rear blade tip.
  • the discharge at the pressure side has the further advantage, besides lower aerodynamic losses, that the cooling air discharged there is in turn sucked into the blade tip gap, so that the external blade tip surface too is supplied with relatively cold air.
  • a protective cover as mentioned is fitted onto a circumferential sealing edge of the blade tip (winglet, squealer) and fastened there.
  • the result is an effective, flat and contourless blade tip geometry with a duct or cavity underneath.
  • the cover is suitably shaped or contoured such that the blade tip can retain the contour of the circumferential sealing edge.
  • the protective cover and/or the cavity on the blade tip can be but do/does not necessarily have to be, designed up to the blade leading edge. Since the pressure level in the front blade tip area is relatively constant and only drops steeply towards the blade trailing edge, it can be advantageous to design the protective cover and hence the duct or cavity only starting from a middle position of the blade tip. With this embodiment, the dust holes can then be near the cover and hence close to the inflow area of the duct or cavity, in order to promote the aspiration of the cold dust hole air into the duct or cavity. It is thus possible to use the dust hole air, otherwise not readily usable for cooling, to cool the pressure-side blade tip close to the trailing edge.
  • the protective cover at the blade leading edge closed (without opening). Hence only one opening of the cavity or of the duct is provided close to the blade trailing edge.
  • the cavity or duct can be flooded completely with cold blade-internal air. This embodiment is suitable in particular for very high hot gas temperatures, when blade-external air is no longer usable for cooling the blade tip.
  • additional webs or supports can be used. By means of these webs, the air can be guided to the thermally most highly loaded areas.
  • the webs or supports are furthermore used as fastening surfaces for the protective cover and thereby contribute to the mechanical stability of the blade tip. Furthermore, the webs or supports can increase the internal heat transfer in the cavity or duct, so that the cooling effect can be further improved.
  • the quantity of film cooling air required for cooling the rotor blade tip can be considerably reduced.
  • FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention
  • FIG. 2 shows a simplified sectional view of a blade tip designed in accordance with the present invention
  • FIG. 3 shows a view, by analogy with FIG. 2 , of a further exemplary embodiment of the present invention
  • FIGS. 4 shows a further exemplary embodiment, by analogy with FIGS. 2 and 3 without lateral blade tip overhang, and
  • FIGS. 5 to 12 show simplified perspective representations of exemplary embodiments in accordance with the present invention.
  • the gas-turbine engine 10 in accordance with FIG. 1 is a generally represented example of a turbomachine where the invention can be used.
  • the engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11 , a fan 12 rotating inside a casing, an intermediate-pressure compressor 13 , a high-pressure compressor 14 , a combustion chamber 15 , a high-pressure turbine 16 , an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19 , all of which being arranged about a central engine axis 1 .
  • the intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20 , generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13 , 14 .
  • the compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17 , respectively.
  • the turbine sections 16 , 17 , 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16 , 17 , 18 , and a subsequent arrangement of turbine rotor blades 24 projecting outwards from a rotatable hub 27 .
  • the compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.
  • FIG. 2 shows a simplified sectional view of a blade tip 29 of a turbine rotor blade 24 .
  • the reference numeral 35 shows the suction side, while the reference numeral 36 indicates the pressure side.
  • Front-side sealing edges 33 are provided on the blade tip 29 .
  • the blade tip 29 can have lateral overhangs (winglets) 34 .
  • a flow duct 37 is provided on the front side of the blade tip 29 and closed by a protective cover 38 .
  • the protective cover 38 is designed profiled, so that any contouring of the blade tip 29 can be retained.
  • FIG. 3 shows a view, by analogy with FIG. 2 , where additional webs or supports 39 are provided which support the protective cover 38 .
  • the webs or supports 39 furthermore enable several flow ducts 37 to be formed, or the airflow through the flow duct 37 to be optimized.
  • FIG. 4 shows a view, by analogy with FIGS. 2 and 3 , where in the exemplary embodiment of FIG. 4 the blade tip has no lateral overhang (winglet overhang).
  • FIGS. 5 to 12 show differing exemplary embodiments of the invention, where the perspective view is schematic and where the protective cover 38 is only shown in simplified form in order to make clear the flow through the flow duct 37 .
  • the front suction-side area of the blade tip 29 has the reference numeral 30
  • the rear area has the reference numeral 31 .
  • a blade trailing edge 32 is designed in the usual way.
  • FIG. 5 shows an exemplary embodiment in which the protective cover 38 is provided on the entire front-side area of the blade tip 29 .
  • the flow necessary for cooling the blade tip is introduced centrally at the blade leading edge, while the outflow through a suction-side opening takes place in the rear area of the blade tip.
  • FIG. 6 shows a design variant of the exemplary embodiment of FIG. 5 , where the protective cover 38 extends over only part of the total length of the blade tip 29 .
  • a centric support 39 is provided in the exemplary embodiment of FIG. 7 , which divides the flow duct 37 .
  • the outflow is provided, in a variation of the exemplary embodiment of FIG. 7 on the pressure side of the rear area 31 of the blade tip 29 .
  • FIG. 9 represents a variant of the exemplary embodiments of FIGS. 7 and 8 and has a centric outlet opening in the area of the blade trailing edge 32 .
  • the protective cover extends completely over the front area of the blade tip 29 and has an outlet opening only on the suction side of the rear area 31 .
  • the cooling air is supplied from the blade interior via ducts.
  • FIG. 11 also shows a protective cover closed in the front area, by analogy with FIG. 10 , where a central web 39 divides the flow duct 37 .
  • FIG. 12 it is provided, in a variation from the exemplary embodiments of FIGS. 10 and 11 , that individual supports acting as turbulators are arranged inside the flow duct 37 .
  • FIGS. 10 to 12 each show the supply of the flow through cooling air holes (dust holes) 41 .
US14/559,668 2013-12-05 2014-12-03 Turbine rotor blade of a gas turbine and method for cooling a blade tip of a turbine rotor blade of a gas turbine Abandoned US20150159488A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102013224998.5 2013-12-05
DE102013224998.5A DE102013224998A1 (de) 2013-12-05 2013-12-05 Turbinenrotorschaufel einer Gasturbine und Verfahren zur Kühlung einer Schaufelspitze einer Turbinenrotorschaufel einer Gasturbine

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EP (1) EP2881541A1 (de)
DE (1) DE102013224998A1 (de)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20140119942A1 (en) * 2012-10-26 2014-05-01 Rolls-Royce Plc Turbine rotor blade of a gas turbine
US20170101889A1 (en) * 2015-10-09 2017-04-13 United Technologies Corporation Multi-flow cooling passage chamber for gas turbine engine
EP3498971A1 (de) * 2017-12-14 2019-06-19 Rolls-Royce plc Tragflügel mit einem teilenblatt
US10344597B2 (en) * 2015-08-17 2019-07-09 United Technologies Corporation Cupped contour for gas turbine engine blade assembly
US10408065B2 (en) 2017-12-06 2019-09-10 General Electric Company Turbine component with rail coolant directing chamber
US10458427B2 (en) * 2014-08-18 2019-10-29 Siemens Aktiengesellschaft Compressor aerofoil
US10570750B2 (en) 2017-12-06 2020-02-25 General Electric Company Turbine component with tip rail cooling passage
US10801334B2 (en) 2018-09-12 2020-10-13 Raytheon Technologies Corporation Cooling arrangement with purge partition
US10844732B2 (en) 2017-12-14 2020-11-24 Rolls-Royce Plc Aerofoil and method of manufacture
US20210239009A1 (en) * 2020-02-04 2021-08-05 United Technologies Corporation Blade with wearable tip-rub-portions above squealer pocket
US20220243597A1 (en) * 2021-02-04 2022-08-04 Doosan Heavy Industries & Construction Co., Ltd. Airfoil with a squealer tip cooling system for a turbine blade, a turbine blade, a turbine blade assembly, a gas turbine and a manufacturing method
CN115126547A (zh) * 2022-05-29 2022-09-30 中国船舶重工集团公司第七0三研究所 一种吸力侧排气的气冷涡轮动叶尾缘结构

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3241991A1 (de) * 2016-05-04 2017-11-08 Siemens Aktiengesellschaft Turbinenanordnung

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6059530A (en) * 1998-12-21 2000-05-09 General Electric Company Twin rib turbine blade
US6422821B1 (en) * 2001-01-09 2002-07-23 General Electric Company Method and apparatus for reducing turbine blade tip temperatures
US20090180895A1 (en) * 2008-01-10 2009-07-16 General Electric Company Turbine blade tip shroud
US7632062B2 (en) * 2004-04-17 2009-12-15 Rolls-Royce Plc Turbine rotor blades
US8672629B2 (en) * 2008-03-05 2014-03-18 Snecma Cooling of the tip of a blade
US8708645B1 (en) * 2011-10-24 2014-04-29 Florida Turbine Technologies, Inc. Turbine rotor blade with multi-vortex tip cooling channels
US8801377B1 (en) * 2011-08-25 2014-08-12 Florida Turbine Technologies, Inc. Turbine blade with tip cooling and sealing

Family Cites Families (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS59231102A (ja) * 1983-06-15 1984-12-25 Toshiba Corp ガスタ−ビンの翼
JPH03194101A (ja) * 1989-12-21 1991-08-23 Toshiba Corp ガスタービン冷却動翼
US5503527A (en) * 1994-12-19 1996-04-02 General Electric Company Turbine blade having tip slot
US7059834B2 (en) * 2003-01-24 2006-06-13 United Technologies Corporation Turbine blade
US6916150B2 (en) * 2003-11-26 2005-07-12 Siemens Westinghouse Power Corporation Cooling system for a tip of a turbine blade
DE502004006484D1 (de) * 2004-01-23 2008-04-24 Siemens Ag Kühlung einer Turbinenschaufel mit einem Doppelboden zwischen Schaufelblatt und Schaufelspitze
GB0724612D0 (en) * 2007-12-19 2008-01-30 Rolls Royce Plc Rotor blades
US8337158B1 (en) * 2009-10-22 2012-12-25 Florida Turbine Technologies, Inc. Turbine blade with tip cap
GB201006451D0 (en) * 2010-04-19 2010-06-02 Rolls Royce Plc Blades
US9546554B2 (en) * 2012-09-27 2017-01-17 Honeywell International Inc. Gas turbine engine components with blade tip cooling

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6059530A (en) * 1998-12-21 2000-05-09 General Electric Company Twin rib turbine blade
US6422821B1 (en) * 2001-01-09 2002-07-23 General Electric Company Method and apparatus for reducing turbine blade tip temperatures
US7632062B2 (en) * 2004-04-17 2009-12-15 Rolls-Royce Plc Turbine rotor blades
US20090180895A1 (en) * 2008-01-10 2009-07-16 General Electric Company Turbine blade tip shroud
US8672629B2 (en) * 2008-03-05 2014-03-18 Snecma Cooling of the tip of a blade
US8801377B1 (en) * 2011-08-25 2014-08-12 Florida Turbine Technologies, Inc. Turbine blade with tip cooling and sealing
US8708645B1 (en) * 2011-10-24 2014-04-29 Florida Turbine Technologies, Inc. Turbine rotor blade with multi-vortex tip cooling channels

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10641107B2 (en) 2012-10-26 2020-05-05 Rolls-Royce Plc Turbine blade with tip overhang along suction side
US9593584B2 (en) * 2012-10-26 2017-03-14 Rolls-Royce Plc Turbine rotor blade of a gas turbine
US20140119942A1 (en) * 2012-10-26 2014-05-01 Rolls-Royce Plc Turbine rotor blade of a gas turbine
US10458427B2 (en) * 2014-08-18 2019-10-29 Siemens Aktiengesellschaft Compressor aerofoil
US10344597B2 (en) * 2015-08-17 2019-07-09 United Technologies Corporation Cupped contour for gas turbine engine blade assembly
US20170101889A1 (en) * 2015-10-09 2017-04-13 United Technologies Corporation Multi-flow cooling passage chamber for gas turbine engine
US10060288B2 (en) * 2015-10-09 2018-08-28 United Technologies Corporation Multi-flow cooling passage chamber for gas turbine engine
US10408065B2 (en) 2017-12-06 2019-09-10 General Electric Company Turbine component with rail coolant directing chamber
US10570750B2 (en) 2017-12-06 2020-02-25 General Electric Company Turbine component with tip rail cooling passage
EP3498971A1 (de) * 2017-12-14 2019-06-19 Rolls-Royce plc Tragflügel mit einem teilenblatt
US10844732B2 (en) 2017-12-14 2020-11-24 Rolls-Royce Plc Aerofoil and method of manufacture
US10968754B2 (en) 2017-12-14 2021-04-06 Rolls-Royce Plc Aerofoil
US10801334B2 (en) 2018-09-12 2020-10-13 Raytheon Technologies Corporation Cooling arrangement with purge partition
US20210239009A1 (en) * 2020-02-04 2021-08-05 United Technologies Corporation Blade with wearable tip-rub-portions above squealer pocket
US11215061B2 (en) * 2020-02-04 2022-01-04 Raytheon Technologies Corporation Blade with wearable tip-rub-portions above squealer pocket
US20220243597A1 (en) * 2021-02-04 2022-08-04 Doosan Heavy Industries & Construction Co., Ltd. Airfoil with a squealer tip cooling system for a turbine blade, a turbine blade, a turbine blade assembly, a gas turbine and a manufacturing method
US11572792B2 (en) * 2021-02-04 2023-02-07 Doosan Enerbility Co., Ltd. Airfoil with a squealer tip cooling system for a turbine blade, a turbine blade, a turbine blade assembly, a gas turbine and a manufacturing method
CN115126547A (zh) * 2022-05-29 2022-09-30 中国船舶重工集团公司第七0三研究所 一种吸力侧排气的气冷涡轮动叶尾缘结构

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DE102013224998A1 (de) 2015-06-11

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