US3700348A - Turbomachinery blade structure - Google Patents

Turbomachinery blade structure Download PDF

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Publication number
US3700348A
US3700348A US752736A US3700348DA US3700348A US 3700348 A US3700348 A US 3700348A US 752736 A US752736 A US 752736A US 3700348D A US3700348D A US 3700348DA US 3700348 A US3700348 A US 3700348A
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US
United States
Prior art keywords
blade
insert
shank
cast member
hollow
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.)
Expired - Lifetime
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US752736A
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English (en)
Inventor
Robert J Corsmeier
Thomas G Howell
Charles E Corrigan
Bernard L Koff
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General Electric Co
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General Electric Co
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Publication date
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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
    • 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
    • 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

  • ABSTRACT An economically fabricated turbomachinery blade structure with good and predictable high temperature characteristics having a hollow integrally cast member defining the blade portion, 'the platform portion, a transition shank portion defining an opening to the blade chamber and a pair of juxtaposed rails extending therefrom.
  • a hollow insert formed with a plurality of nozzles in the side walls thereof, is secured within the cast member chamber so as to direct coolant fluid against the inner surfaces of the blade portion.
  • the juxtaposed rails are formed into abutting contact at their free ends and bonded.
  • a root insert is provided which is secured to the inner surfaces of shank rails.
  • Suitable passages are provided to direct coolant fluid into the insert chamber and to discharge the coolant fluid through the trailing edge. and the end closure of the blade.
  • Means may be provided to automatically space the insert nozzles from the blade side walls.
  • a plurality of posts may be provided in the trailing edge portion of the blade to increase the effective convection cooling area. All bonds subjected to structural loading are spaced from the blade portion and positioned to be bathed in coolant fluid so as to reduce their operating temperature environment.
  • Such prior blade structures have usually been complex in nature and require sophisticated and expensive machining and assembly techniques in their fabrication.
  • many such prior blade structures employ one or more brazed, welded, or similarly bonded joints in the portion of the blade structure exposed to the hot gas stream.
  • bonds may impose a severe limitation on the operating temperature capacity of the blade and such limitation will be greatly enhanced is such bond is operative to transmit blade loading.
  • This invention is concerned with a fluid cooled, high temperature blade structure which overcomes theabove mentioned problems.
  • a primary object of this invention is a relatively low cost, high temperature blade structure having all bonds spaced from the portion of the blade structure in contact with the hot gas stream.
  • Another object of this invention is a turbine blade structure having predictable high temperature characteristics.
  • the turbomachinery blade structure of this invention comprises a cast member, means for mounting the cast member to a supporting rotor, a insert for the cast member formed with a chamber therein, and means for cooling said blade structure by impingement of a cooling fluid on the inner surfaces of the blade portion of the cast member.
  • the cooling fluid is discharged into the hot gas stream through trailing passages. All structural bonds between the cast member, insert, and mounting means are spaced from the hot gas stream.
  • FIG. 1 is a partial cross-sectional view diagrammatically showing one installation of the blade structure of this invention
  • FIG. 2 is an exploded perspective view, in partial section, showing one embodiment of the blade structure of this invention
  • FIG. 3 is a cross-sectional view "taken along line 33 of FIG. 2;
  • FIG. 4 is a cross-sectional view taken along line 4-4 of FIG. 2;
  • FIG. 5 is a partial cross-sectional view taken along line 5-5 of FIG. 2;
  • FIG. 6 is a cross-sectional view, like that of FIG. 3, of another embodiment of the blade structure of this inventron;
  • FIG. 7 is a cross-sectional view taken along line 7-7 of FIG. 6;
  • FIG. 8 is a cross-sectional view taken along line 8-8 of FIG. 6.
  • the turbine portion of a gas turbine engine has been shown diagrammatically in partial section as comprising a rotor 10 having a plurality of blades 12 extending radially therefrom around its periphery.
  • the blades 12 extend into a passage 14 through which a motive fluid or hot gas stream (indicated by flow arrows in FIG. 1) is directed and are adapted to extract energy from the motive fluid and impart rotary motion to the rotor 10 in a well known manner.
  • suitable means shown generally at 16 may be provided to direct a coolant fluid (indicated generally by the broken flow arrows in FIG. I) through the rotor 10 to each blade 12.
  • FIGS. 2 through 5 one embodiment of the improved blade structure of this invention has been shown as including a hollow cast member 20, means 22 for engaging and securing the blade to the supporting rotor 10, and an insert .24 for directing coolzaart fluid within a blade portion 26 of the cast member
  • the cast member has been shown as including a platform portion 28, a blade portion 26 extending or projecting from the platform portion 28, and a transition shank portion 30 extending from the platform portion 28, generally oppositely to the blade portion 26, and defining an opening or passage 32 to a chamber 34 within the blade portion 26.
  • the blade portion 26 is formed with a leading edge 36, a trailing edge 38 and airfoil shaped side walls 40 extending therebetween which are suitably formed and adapted to extract energy from the motive fluid flowing thereacross, as in FIG. I, through the mechanism of either action or reaction fluid dynamics and impart rotary motion to the supporting rotor 10.
  • the chamber 34 is defined by blade side wall inner surfaces 42 and an integrally east end closure wall 44.
  • the cast member chamber 34 includes a railing edge portion 46 having a plurality of posts 48 extending thereacross and connecting opposed side walls 40.
  • a plurality of passages 50 are provided in trailing edge 38 to enable efflux of coolant fluid from chamber 34 as will be hereinafter discussed.
  • Additional efflux passages 52 may be formed in end closure 44 to provide cooling to that area of the blade.
  • the platform portion 28 has been shown as having a top surface 54 which is adapted to define, in part, the passage 14 into which the blade portion 26 extends and through which the motive fluid or hot gas stream is directed.
  • the platform portion may be variously shaped to perform interlocking or dampening functions between adjacent blades when arranged and mounted in a circumferential row around the supporting rotor 10 as in FIG. 1.
  • case member shank portion 30 is integrally cast with and extends from platform portion 28, generally oppositely of the blade portion 26, and defines the opening or passage 32 communicating with cast member chamber 34.
  • FIGS. 2 through the means 22 for attaching the blade to the rotor has been shown as comprising juxtaposed shank rails 56, integrally cast with member 20, which extend from the shank 30 generally oppositely to blade portion 26.
  • the rails 56 are suitably formed into abutting contact and suitably bonded together as at 58.
  • the outer surfaces of the joined shank rails 56 are suitably formed in a fir tree, dovetail, or like configuration, as shown generally at 60, so as to enable attachment of the blade structure to the supporting rotor 10 in a well known manner.
  • a plurality of passages 62 are provided through the rail abutment to enable influx of a coolant fluid to the blade structure as will be hereinafter discussed.
  • the insert 24 has been shown as being a thin walled shell having side walls 64 generally conforming to the contour of the blade side walls 40 and formed with a plurality of nozzles or apertures 66 which are disposed in spaced relationship to the inner surfaces 42 of the blade portion 26 by suitable spacing means 68. As viewed in FIG. 2, the insert 24 is closed at one end 70 and defines a chamber 72 therein which is open at end 74 to receive coolant fluid.
  • Suitable means as at 76 are provided to secure the insert 24 to the cast member 20 in a manner enabling passage of coolant fluid into chamber 72.
  • such means also function to close passage 32 and comprise a collar 78 formed with a passage 80 and a surrounding flange 82.
  • the collar 78 is bonded to the open end 74 of insert 24 with passage 80 communicating with chamber 72 and the flange 82 abuts and is suitably bonded to a lower rim 84 of the transition shank 30 in a manner securing the insert 24 to the cast member 20 and closing passage 32.
  • seal plates 86 are provided in order to direct coolant fluid from passages 62 to collar passage 80 and, hence, into the insert chamber 72.
  • suitable means such as seal plates 86 are provided.
  • one seal plate extends between rail leading edges 88 in overlapping relationship with shank 30 and the rail abutment while the other seal plate similarly extends between rail trailing edges 90.
  • spacing means 68 has been shown as a plurality of dimples, projections or the like, integrally formed with insert side walls 64 and abutting inner surfaces 42 of the blade side walls 40, it should be understood that such projections may be carried by the cast member 20.
  • the blade attachment means 22 has been shown as comprising a root insert 92 which abuts and is suitably bonded to the opposed inner shank rail surfaces 94 and 96.
  • the insert 24 is secured at its open end 74 with chamber 72 communicating with passages 62.
  • the root insert 92 is suitably formed with a fir tree or other suitable configuration, as at 60 and passages 62 are provided through the mounting means for delivery of coolant fluid into the blade structure.
  • means such as the seal plates 86 may be provided as in the embodiment of FIG. 2.
  • a primary advantage of the turbomachinery blade structure of this invention is the economical manner in which it may be fabricated.
  • the cast member 20, which includes the blade portion 26, platform portion 28, transition shank 30 and rails 56 is formed as a relatively non complex integral casting.
  • the lower ends of shank rails 56 are cast in a spaced apart, flared out position, as at 98 in FIG. 5, so as to allow insertion therebetween of the thin walled insert 24.
  • this assembly is simply inserted into cast member chamber 34, with means 68 automatically establishing the proper spacing between the insert nozzles 66 and the blade inner surface 42.
  • the ends of the shank rails 56 are suitably formed from their flared out position 98 into abutting contact.
  • the shank rails are then bonded at 58 and the proper external contour, as at 60, is machined.
  • the passages 62 are preferably cast but may be machined after bonding.
  • the insert 24 may be economically formed of any suitable metallic material by a deep drawing, stamping or like suitable process, with the closed end being either integrally formed with side walls 64 or bonded closed.
  • FIGS. 2 through 5 is also extremely advantageous from the structural standpoint in that in transmitting either centrifugal or aerodynamicloading to the support rotor 10 it is not necessary to transmit force through a bonded joint. It should also be noted that no bond is directly exposed to the high temperature environment of the motive fluid and that all bonds are arranged so as to be continually bathed in coolant fluid.
  • the cast member 20 is formed as an integral casting and is conveniently assembled with the insert 24, after the latter has been secured to root insert 92, by simply sliding insert 24 within the chamber 34 and bonding the root insert 92 to inner rail surfaces 94 and 96.
  • the embodiment of FIG. 6 enables the material selection of the blade portion 26 and the mounting means 22 to be tailored to their respective environments, it being understood that the stress loading and temperature environment of the blade portion 26 is somewhat different from the temperature environment and structural loading of the mounting means 22.
  • the blade structure of this invention is mounted as one of a circumferential row of blades in a well known manner to the supporting rotor by engaging the mounting means 22 with a correspondingly contoured groove (not shown) in the rotor 10.
  • the blade portion 26 extends generally radially from the supporting rotor and projects into the annular passage 14 through which the motive fluid is directed.
  • the blade portion 26, through the well known mechanics of action and/or reaction fluid dynamics, extracts energy from the hot gas stream and imparts rotary motion to the supporting rotor 10. In performing this function, it is well known that the efficiency, in terms of thrust or useful work output per unit of fuel input, is greatly increased by increasing the turbine inlet temperature of the motive fluid.
  • suitable means 16 are provided to direct coolant fluid to inlet passage 62. From inlet passages 62, the coolant fluid enters insert chamber 72 and is then impinged against the blade wall inner surfaces 42 by nozzles 66. As will be understood, the number, location and size of the nozzles 66 may be varied so as to achieve a generally uniform side wall temperature around the periphery of the blade.
  • the coolant fluid is exhausted into the hot gas stream through end closure passages 52 and trailing edge passages 50.
  • Posts 48 are provided in chamber portion 46 to increase the convection heat transfer to the fluid passing into the hot gas stream through trailing edge passages 50.
  • the blade structure of this invention does not rely on transpiration, film or other like modes of cooling which may be affected by the flow patterns of the motive fluid around the blade, cooling effectivity and, therefore, the temperature capability of the blade is highly predictable.
  • the present invention provides an economically fabricated, non complex cooled blade structure which possesses good and predictable high temperature characteristics.
  • a turbomachinery blade structure comprising:
  • a hollow, integrally cast member including a platform portion adapted to define, in part, an inner boundary of a motive fluid passage, a blade portion projecting from said platform portion, a transition shank projecting from said platform portion generally oppositely to said blade portion and defining an opening to said hollow cast member, and a pair of juxtaposed shank rails projecting from said transition shank;
  • a hollow insert slideably received within said hollow cast member blade portion in spaced relationship thereto, said insert formed with an open inner end for receiving a coolant fluid, a closed outer end and a plurality of apertures forimpinging said coolant fluid against said cast member blade portion;
  • a mounting collar formed with a passage therethrough, the open end of said insert secured to said collar in flow communication with said collar passage, said collar including a flange secured to said tubular shank in closing relationship to said shank opening;
  • attachment means enabling attachment of said blade structure to a supporting rotor with said blade portion extending generally radially therefrom, said attachment means comprising abutting, bonded together radial inner ends of said shank rails;
  • said blade portion includes a leading edge, a trailing edge, generally airfoil shaped. sidewalls extending therebetween, and an end closure wall, said means for discharging coolant fluid from said cast member chamber comprising a plurality of passages formed through said trailing edge and said end closure wall.
  • a turbomachinery blade structure comprising:
  • a hollow, integrally cast member including a platform portion adapted to define, in part, an inner boundary of a motive fluid passage, a blade portion projecting from said platform portion, a transition shank portion projecting from said platform portion generally oppositely to said blade portion and defining an opening to said hollow cast member, and a pair of juxtaposed shank rails projecting from said transition shank;
  • a hollow insert slideably received within said hollow cast member blade portion in spaced relationship thereto, said insert formed with an open inner end for receiving a coolant fluid, a closed outer end and a plurality of apertures for impinging said coolant fluid against said cast member blade portion;
  • said attachment means comprising a root insert extending between and bonded to said shank rails, and said root insert formed with at least one radially extending passage therethrough, said hollow insert secured to said root insert with said hollow insert open end in flow communication with said root insert passage;

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US752736A 1968-08-13 1968-08-13 Turbomachinery blade structure Expired - Lifetime US3700348A (en)

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Cited By (39)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3799696A (en) * 1971-07-02 1974-03-26 Rolls Royce Cooled vane or blade for a gas turbine engine
US3809494A (en) * 1971-06-30 1974-05-07 Rolls Royce 1971 Ltd Vane or blade for a gas turbine engine
US3819295A (en) * 1972-09-21 1974-06-25 Gen Electric Cooling slot for airfoil blade
US3834831A (en) * 1973-01-23 1974-09-10 Westinghouse Electric Corp Blade shank cooling arrangement
US3846041A (en) * 1972-10-31 1974-11-05 Avco Corp Impingement cooled turbine blades and method of making same
US3864199A (en) * 1973-07-26 1975-02-04 Gen Motors Corp Angular discharge porous sheet
US3866448A (en) * 1973-01-02 1975-02-18 Gen Electric Apparatus for constructing air cooled turbomachinery blading
US3867068A (en) * 1973-03-30 1975-02-18 Gen Electric Turbomachinery blade cooling insert retainers
JPS5048309A (fr) * 1973-08-31 1975-04-30
US3902820A (en) * 1973-07-02 1975-09-02 Westinghouse Electric Corp Fluid cooled turbine rotor blade
US3921271A (en) * 1973-01-02 1975-11-25 Gen Electric Air-cooled turbine blade and method of making same
US3930748A (en) * 1972-08-02 1976-01-06 Rolls-Royce (1971) Limited Hollow cooled vane or blade for a gas turbine engine
US3934322A (en) * 1972-09-21 1976-01-27 General Electric Company Method for forming cooling slot in airfoil blades
US3966357A (en) * 1974-09-25 1976-06-29 General Electric Company Blade baffle damper
US3973874A (en) * 1974-09-25 1976-08-10 General Electric Company Impingement baffle collars
US3982854A (en) * 1971-12-20 1976-09-28 General Electric Company Friction welded metallic turbomachinery blade element
US3994622A (en) * 1975-11-24 1976-11-30 United Technologies Corporation Coolable turbine blade
FR2335807A1 (fr) * 1975-12-20 1977-07-15 Rolls Royce Dispositif de refroidissement d'une surface par impact d'un fluide refrigerant
US4183716A (en) * 1977-01-20 1980-01-15 The Director of National Aerospace Laboratory of Science and Technology Agency, Toshio Kawasaki Air-cooled turbine blade
US4359310A (en) * 1979-12-12 1982-11-16 Bbc Brown, Boveri & Company Limited Cooled wall
US5104290A (en) * 1989-11-09 1992-04-14 Rolls-Royce Plc Bladed rotor with axially extending radially re-entrant features
EP0974735A2 (fr) 1998-07-20 2000-01-26 General Electric Company Chicane avec des protubérances
US6193465B1 (en) * 1998-09-28 2001-02-27 General Electric Company Trapped insert turbine airfoil
FR2798422A1 (fr) * 1990-01-24 2001-03-16 United Technologies Corp Pales refroidies pour moteur a turbine a gaz
US6238182B1 (en) 1999-02-19 2001-05-29 Meyer Tool, Inc. Joint for a turbine component
US6742984B1 (en) * 2003-05-19 2004-06-01 General Electric Company Divided insert for steam cooled nozzles and method for supporting and separating divided insert
US7695245B1 (en) * 2007-03-06 2010-04-13 Florida Turbine Technologies, Inc. Turbine airfoil with a multi-impingement cooled spar and shell
JP2010518300A (ja) * 2007-02-01 2010-05-27 シーメンス アクチエンゲゼルシヤフト タービン翼
US8449249B2 (en) 2010-04-09 2013-05-28 Williams International Co., L.L.C. Turbine nozzle apparatus and associated method of manufacture
US20130202409A1 (en) * 2010-04-29 2013-08-08 Richard Jones Turbine vane hollow inner rail
US8622702B1 (en) 2010-04-21 2014-01-07 Florida Turbine Technologies, Inc. Turbine blade with cooling air inlet holes
JP2014177943A (ja) * 2013-03-14 2014-09-25 General Electric Co <Ge> ガスタービンエンジンのタービン動翼の冷却通路
US20150198050A1 (en) * 2014-01-15 2015-07-16 Siemens Energy, Inc. Internal cooling system with corrugated insert forming nearwall cooling channels for airfoil usable in a gas turbine engine
US20160222796A1 (en) * 2013-09-18 2016-08-04 United Technologies Corporation Manufacturing method for a baffle-containing blade
CN107416184A (zh) * 2017-08-01 2017-12-01 西华酷农无人机产业园运营有限公司 一种高效散热的无人机螺旋桨
US10612397B2 (en) * 2016-02-22 2020-04-07 Mitsubishi Hitachi Power Systems, Ltd. Insert assembly, airfoil, gas turbine, and airfoil manufacturing method
US10738622B2 (en) 2016-08-09 2020-08-11 General Electric Company Components having outer wall recesses for impingement cooling
US11149568B2 (en) * 2018-12-20 2021-10-19 Rolls-Royce Plc Sliding ceramic matrix composite vane assembly for gas turbine engines
US11346246B2 (en) * 2017-12-01 2022-05-31 Siemens Energy, Inc. Brazed in heat transfer feature for cooled turbine components

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB685769A (en) * 1949-11-22 1953-01-14 Rolls Royce Improvements relating to compressor and turbine blading
US2647368A (en) * 1949-05-09 1953-08-04 Hermann Oestrich Method and apparatus for internally cooling gas turbine blades with air, fuel, and water
DE926397C (de) * 1952-09-25 1955-04-18 Ernst Heinkel Dr Dr Hohlschaufel, vorzugsweise fuer axial durchstroemte Turbinen
US2979809A (en) * 1956-03-14 1961-04-18 Napier & Son Ltd Method of making hollow turbine blades
US3032314A (en) * 1957-05-28 1962-05-01 Snecma Method of and device for cooling the component elements of machines

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2647368A (en) * 1949-05-09 1953-08-04 Hermann Oestrich Method and apparatus for internally cooling gas turbine blades with air, fuel, and water
GB685769A (en) * 1949-11-22 1953-01-14 Rolls Royce Improvements relating to compressor and turbine blading
DE926397C (de) * 1952-09-25 1955-04-18 Ernst Heinkel Dr Dr Hohlschaufel, vorzugsweise fuer axial durchstroemte Turbinen
US2979809A (en) * 1956-03-14 1961-04-18 Napier & Son Ltd Method of making hollow turbine blades
US3032314A (en) * 1957-05-28 1962-05-01 Snecma Method of and device for cooling the component elements of machines

Cited By (46)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3809494A (en) * 1971-06-30 1974-05-07 Rolls Royce 1971 Ltd Vane or blade for a gas turbine engine
US3799696A (en) * 1971-07-02 1974-03-26 Rolls Royce Cooled vane or blade for a gas turbine engine
US3982854A (en) * 1971-12-20 1976-09-28 General Electric Company Friction welded metallic turbomachinery blade element
US3930748A (en) * 1972-08-02 1976-01-06 Rolls-Royce (1971) Limited Hollow cooled vane or blade for a gas turbine engine
US3819295A (en) * 1972-09-21 1974-06-25 Gen Electric Cooling slot for airfoil blade
US3934322A (en) * 1972-09-21 1976-01-27 General Electric Company Method for forming cooling slot in airfoil blades
US3846041A (en) * 1972-10-31 1974-11-05 Avco Corp Impingement cooled turbine blades and method of making same
US3921271A (en) * 1973-01-02 1975-11-25 Gen Electric Air-cooled turbine blade and method of making same
US3866448A (en) * 1973-01-02 1975-02-18 Gen Electric Apparatus for constructing air cooled turbomachinery blading
US3834831A (en) * 1973-01-23 1974-09-10 Westinghouse Electric Corp Blade shank cooling arrangement
US3867068A (en) * 1973-03-30 1975-02-18 Gen Electric Turbomachinery blade cooling insert retainers
US3902820A (en) * 1973-07-02 1975-09-02 Westinghouse Electric Corp Fluid cooled turbine rotor blade
US3864199A (en) * 1973-07-26 1975-02-04 Gen Motors Corp Angular discharge porous sheet
JPS5048309A (fr) * 1973-08-31 1975-04-30
US3973874A (en) * 1974-09-25 1976-08-10 General Electric Company Impingement baffle collars
US3966357A (en) * 1974-09-25 1976-06-29 General Electric Company Blade baffle damper
US3994622A (en) * 1975-11-24 1976-11-30 United Technologies Corporation Coolable turbine blade
FR2335807A1 (fr) * 1975-12-20 1977-07-15 Rolls Royce Dispositif de refroidissement d'une surface par impact d'un fluide refrigerant
US4105364A (en) * 1975-12-20 1978-08-08 Rolls-Royce Limited Vane for a gas turbine engine having means for impingement cooling thereof
US4183716A (en) * 1977-01-20 1980-01-15 The Director of National Aerospace Laboratory of Science and Technology Agency, Toshio Kawasaki Air-cooled turbine blade
US4359310A (en) * 1979-12-12 1982-11-16 Bbc Brown, Boveri & Company Limited Cooled wall
US5104290A (en) * 1989-11-09 1992-04-14 Rolls-Royce Plc Bladed rotor with axially extending radially re-entrant features
FR2798422A1 (fr) * 1990-01-24 2001-03-16 United Technologies Corp Pales refroidies pour moteur a turbine a gaz
EP0974735A2 (fr) 1998-07-20 2000-01-26 General Electric Company Chicane avec des protubérances
JP2000130760A (ja) * 1998-07-20 2000-05-12 General Electric Co <Ge> ディンプルを設けたインピンジメントバッフル
EP0974735A3 (fr) * 1998-07-20 2001-05-16 General Electric Company Chicane avec des protubérances
US6193465B1 (en) * 1998-09-28 2001-02-27 General Electric Company Trapped insert turbine airfoil
US6238182B1 (en) 1999-02-19 2001-05-29 Meyer Tool, Inc. Joint for a turbine component
US6742984B1 (en) * 2003-05-19 2004-06-01 General Electric Company Divided insert for steam cooled nozzles and method for supporting and separating divided insert
JP2010518300A (ja) * 2007-02-01 2010-05-27 シーメンス アクチエンゲゼルシヤフト タービン翼
US7695245B1 (en) * 2007-03-06 2010-04-13 Florida Turbine Technologies, Inc. Turbine airfoil with a multi-impingement cooled spar and shell
US8449249B2 (en) 2010-04-09 2013-05-28 Williams International Co., L.L.C. Turbine nozzle apparatus and associated method of manufacture
US8622702B1 (en) 2010-04-21 2014-01-07 Florida Turbine Technologies, Inc. Turbine blade with cooling air inlet holes
US9869200B2 (en) * 2010-04-29 2018-01-16 Siemens Aktiengesellschaft Turbine vane hollow inner rail
US20130202409A1 (en) * 2010-04-29 2013-08-08 Richard Jones Turbine vane hollow inner rail
JP2014177943A (ja) * 2013-03-14 2014-09-25 General Electric Co <Ge> ガスタービンエンジンのタービン動翼の冷却通路
US9567859B2 (en) * 2013-03-14 2017-02-14 General Electric Company Cooling passages for turbine buckets of a gas turbine engine
US20140286771A1 (en) * 2013-03-14 2014-09-25 General Electric Company Cooling passages for turbine buckets of a gas turbine engine
US20160222796A1 (en) * 2013-09-18 2016-08-04 United Technologies Corporation Manufacturing method for a baffle-containing blade
EP3049625A4 (fr) * 2013-09-18 2017-07-19 United Technologies Corporation Procédé de fabrication pour une pale contenant un déflecteur
US20150198050A1 (en) * 2014-01-15 2015-07-16 Siemens Energy, Inc. Internal cooling system with corrugated insert forming nearwall cooling channels for airfoil usable in a gas turbine engine
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