US3867068A - Turbomachinery blade cooling insert retainers - Google Patents

Turbomachinery blade cooling insert retainers Download PDF

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Publication number
US3867068A
US3867068A US346422A US34642273A US3867068A US 3867068 A US3867068 A US 3867068A US 346422 A US346422 A US 346422A US 34642273 A US34642273 A US 34642273A US 3867068 A US3867068 A US 3867068A
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US
United States
Prior art keywords
insert
blade
bushing
flange
pin
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
Application number
US346422A
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English (en)
Inventor
Robert J Corsmeier
Charles E Corrigan
Ronald E Dennis
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
Priority to US346422A priority Critical patent/US3867068A/en
Priority to GB1120974A priority patent/GB1467374A/en
Priority to CA195,247A priority patent/CA994672A/en
Priority to DE2413292A priority patent/DE2413292A1/de
Priority to IT49539/74A priority patent/IT1003841B/it
Priority to SE7404215A priority patent/SE390433B/xx
Priority to NL7404257A priority patent/NL7404257A/xx
Priority to JP49034775A priority patent/JPS5026103A/ja
Priority to BE142677A priority patent/BE813090A/xx
Priority to FR7411191A priority patent/FR2223550B1/fr
Application granted granted Critical
Publication of US3867068A publication Critical patent/US3867068A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—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

  • An improved turbomachinery blade includes impingement inserts adapted to direct a coolant toward surfaces defining an internal cavity within the blade.
  • the inserts are provided with means which mechanically secure the inserts in place within the cavity.
  • the securing means include a bushing having flanges of approximately the same thickness as that of the side walls of the insert, which flanges are secured within openings provided in such side walls near the bottom end of the inserts.
  • a pin is inserted within holes provided in the root portion of the blade and a hole extending through the bushing, which is aligned with the holes in the root portion of the blade, and a collar of braze alloy is positioned around each end of the pin, thereby fluidically sealing each end of the pin and the insert and securing the pin to the blade and the insert in place within the cavity.
  • This invention relates generally to gas turbine engines and, more particularly, to an improved fluid cooled turbomachinery blade structure for use in high temperature gas turbines.
  • Coolant is delivered to the interior of such an insert and is expelled through a multiplicity of small holes against an internal wall of the turbomachinery blade, thereby cooling the portion of the turbomachinery blade which is exposed to the hot gas stream.
  • the overall weight of a gas turbine engine, and in particular the thrust-to-weight ratio of the engine is one of the engines most critical characteristics.
  • the weight of the turbine rotor assembly like that of any other turbomachinery component, must thereforebe maintained at the minimum practical levels obtainable.
  • the wall thickness of the hollow body portion of the blade itself not only must the wall thickness of the hollow body portion of the blade itself be maintained at a minimum thickness, but the wall thickness of the impingement insert also must be maintained at minimum practical levels. Wall thicknesses for such inserts on the order of 0.010 inch are becoming more and more prevalent.
  • One method of keeping the blade weight low is to make the airfoil walls as thin as possible.
  • a minimum practical wall thickness is chosen for the tip portion of the blade, but the airfoil wall thickness must increase as it nears the blade platform in order to carry the increasing load of the airfoil.
  • This centrifugal loading and the radial temperature gradient of the airfoil combine to produce blades having walls that are constant in thickness from the tip to a point above the pitch line, thereafter increase in thickness to a short distance below the pitch line, and then become essentially constant in thickness to the airfoil root portion.
  • the internal cavity formed by the hollow body portion of the blade also changes in size, with a blade as described above providing a bottle-shaped cavity with the largest opening being near the tip of the blade.
  • an insert which includes an opening which is drilled or punched through both sides thereof near the bottom end thereof.
  • a grommettype bushing is placed in the openings and secured to both sides of the insert.
  • a hole is then drilled through the bushing and the insert is pushed into the blade internal cavity until the hole in the bushing lines up with suitable holes in the shank portion of the blade.
  • a pin is then inserted into the aligned holes and braze alloy is placed around each end of the pin and the insert base in order to seal the ends of the pin and the insert base and retain the pin in place.
  • the bushing is provided with flanges of approximately the same thickness as that of the walls of the insert so as to lend itself to an optimum weldment.
  • FIG. 1 is a partial, cross-sectional view diagrammatically showing one installation of the inventive insert retaining means of the present invention
  • FIG. 2 is a cross-sectional view taken along line 22 of FIG. 1; 7
  • FIG. 3 is a cross-sectional view taken along line 33 of FIG. 1;
  • FIG. 4 is an enlarged, partial, cross-sectional view showing the details of a portion of FIG. 1;
  • FIG. 5 is a partial, sectional view of an alternative retaining means
  • FIG. 6 is a cross-sectional view taken generally along line 66 of FIG. 5;
  • FIG. 7 is an axial cross-sectional view of a second alternative retaining means
  • FIG. 8 is a cross-sectional view, with portion deleted, taken generally along line 8-8 of FIG. 7;
  • FIG. 9 is a partial sectional view, similar to FIG. 5, of another alternative embodiment.
  • FIG. 10 is a cross-sectional view, with portions deleted, taken generally along line 10-10 of FIG. 9.
  • FIGS. 1 through 4 wherein a turbomachinery blade incorporating the present inventive insert retaining means is generally designated by the numeral 10.
  • the blade 10 includes a root portion 12 which provides a pair of tangs 14 adapted to mount within a dovetail slot (not shown) associated with a gas turbine rotor disc (not shown).
  • the blade 10 further includes an airfoil-shaped, hollow body portion 16 which is separated from the root portion 12 by means of a blade platform 18.
  • the airfoil-shaped, hollow body portion 16 includes a leading edge 20 and a trailing edge 22 and a pair of airfoil-shaped side walls 24 and 26 extending therebetween which are suitably formed and adapted to extract energy from a motive fluid flowing thereacross.
  • the side walls 24 and 26 cooperate to define an internal cavity 28 which, in the present instance, is divided into two separate sections, leading edge cavity 30 and trailing edge cavity 32 by means of a rib member 34, which is formed integrally with hollow body portion 16 and extends between the side walls 24 and 26.
  • the side walls 24 and 26 vary in thickness from a relatively thin cross section near the tip of the hollow body portion 16 to a thicker cross section near the blade platform 18.
  • the leading edge cavity 30 and the trailing edge cavity 32 are bottle-shaped cavities which are wider near the tip end of the hollow body portion 16 than near the blade platform 18.
  • a thin sheet metal impingement insert 36 is positioned within the leading edge cavity 30, while a similar impingement insert 38 is located within the trailing edge cavity 32.
  • Each of the inserts 36 and 38 is shown as being a thin-walled shell having side walls 40 and 42 generally conforming to the blade side walls 24 and 26.
  • the side walls 40 and 42 are formed with a plurality of nozzles or apertures 44 which are disposed in spaced relationship to the inner surfaces of the cavities 30 and 32 by suitable spacing means 46.
  • the inserts 36 and 38 are closed at one end 48 and define a chamber 50 therein which is open at end 52 to receive coolant fluid.
  • suitable passage means 54 are provided in the root portion 12 of the blade 10. Furthermore, the open or base end 52 of the inserts 36 and 38 are sized so as to sealingly engage side walls 56 located near the top of the passage means 54. As described in greater detail hereafter, the joint b etween the base end 52 and the side walls 56 can be filled with a braze alloy to further seal this joint. In this manner. coolant delivered through the passage means 54 is directed to the chamber 50 and thus to the plurality of apertures or nozzles 44 and is not permitted to pass around the open ends 52 of the inserts 36 and 38 directly into the cavities 30 and 32.
  • the coolant fluid in passing through the apertures 44, impinges directly on the inner sides of the leading edge 20 and the side walls 24 and 26 of the blade 10, thereby cooling such walls in a known manner.
  • the coolant fluid in the leading edge cavity 30 then exits through openings 60 (one of which is shown in FIG. 1) provided in a tip cap 62, which defines the outer end of the cavities 30 and 32, and through film openings (not shown) in the hollow body portion 16, while the coolant fluid in the trailing edge cavity 32 exits through openings in the tip cap and through trailing edge slots 64 located in the trailing edge 22 of the blade 10.
  • means are provided for mechanically securing the inserts 36 and 38 into the position shown in FIG. 1.
  • the means for mechanically securing the inserts in place are generally designated by the numeral 66 and include a grommettype bushing 68 which is positioned within openings 70 and 72 located in the side walls 40 and 42 of the inserts 36 and 38. Since the means for mechanically securing the inserts 36 and 38 in place are substantially identical, only the structure associated with the insert 36 will be described.
  • the bushing 68 includes a pair of thin flanges 74 and 76 which are interconnected by means of an inner spool 78.
  • the flanges 74 and 76 are sized so as to correspond in shape to that of the openings 70 and 72 associated with the side walls 40 and 42.
  • the openings 70 and 72 are made to an approximately circular configuration, although other configurations would also be suitable.
  • the flanges 74 and 76 are made of approximately the same thickness as that of the side walls 40 and 42 so as to provide a configuration which lends itself to an optimum weldment.
  • the openings 70 and 72 provide a much larger perimeter than that of the inner spool 78 so as to provide a large weld area which further enhances the strength of the weldment.
  • the inner spool 78 is provided with a pair of flats 80 and 82 which extend parallel to the side walls 56 of the passage means 54 when the flanges 74 and 76 are positioned within the openings 70 and 72.
  • the flats 80 and 82 permit the correct volume of air to pass through the open end 52 and into the chamber 50 of the insert 36.
  • the bushing 68 Prior to assembling the insert 36 into the blade 10, the bushing 68 is positioned within the insert and is welded in place. A hole 84 is then drilled through the center of the bushing 68, and the insert 36 is positioned within the hollow body portion 16 of the blade through the open end thereof. The hole 84 is aligned with a pair of holes 86 and 88 located within the root portion 12 of the blade 10. A pin 90 is then inserted through the holes 84, 86 and 88, as shown in FIG. 3.
  • the holes 86 and 88 in the root portion of the blade 10 are made oversized with respect to the diameter of the pin 90 so as to allow for the tolerance build-up of the various parts and also the placement of a braze collar around the ends of the pin 90.
  • the hole 84 is sized so as to fit relatively securely around thepin 90 was to provide accurate placement of the insert 36 within the hollow body portion 16.
  • the braze alloy collar which is placed around both ends of the pin 90 not only secures the pin 90 to the root portion 12 of the blade but also acts as a seal which precludes the flow of coolant around the pin 90 before it enters the chamber 50 associated with the insert 36.
  • the base end 52 of the insert can be sized so as to provide a slight gap between the outer wall of the insert 36 and the side walls 56 of the blade.
  • the braze not only seals the pins 90 but also flows into the gap between the insert and the wall 56 and effectively seals this gap. In this manner, all coolant which flows through the passage means 54 must flow into the chamber 50 associated with one of the inserts 36 or 38.
  • the braze alloy is placed around each of the pins 90.
  • the tip cap 62 is then located in place, a suitable braze alloy is applied to the tip cap 62, and the entire blade is then heated which causes the braze alloy to melt, thereby sealing each end of the pins 90 and the base ends of the inserts in addition to securing the pins 90 and the tip cap 62 in their respective positions.
  • FIGS. 5 through 10 a number of alternative embodiments of the mechanical securing means for cooling inserts are shown.
  • a bushing 94 having a pair of relatively elongated flanges 96 and 98 could be utilized in place of the bushing 68.
  • the flanges could extend down to the open end of the insert.
  • An inner spool 100 interconnects the flanges 96 and 98, and, if desired, flats 102 and 104 can be provided on either side of the spool 100.
  • a hole 106 is drilled therethrough and a pin 108 is positioned therein in such a manner as to secure insert 36' to the blade 10.
  • a bushing 110 is designed to fit within the open end 52 of the insert 36.
  • the bushing 110 would have an outer contour shaped in the form of an airfoil and sized so as to fit within an insert 36".
  • the bushing 110 would be hollow and would provide a passageway 112 for the flow of coolant into a chamber 50" of the insert 36''.
  • the bushing 110 In assembling the insert 36 into the blade, the bushing 110 could be initially brazed in place within the insert 36", and the insert 36" thereafter positioned within the blade 10. Holes 114 and 116 could then be drilled through the root portion 12 and each side of the insert 36. Pins 118 and 120 would then be positioned within the holes 114 and 116 and would act to secure the bushing and, thus, the insert 36" in place.
  • bushing 68 described in connection with FIGS. 1 through 4 provides an extremely secure assembly
  • a simple cylindrical collar or a tear-shaped bushing such as that shown at in FIGS. 9 and 10 may be all that is required to secure the insert in place.
  • the bushing 130 would be welded or brazed to the inner sides of the insert 36" and a hole 132 would be drilled through the insert 36" and bushing 130.
  • a pin 134 would be inserted within the hole 132 and brazed in place as described above.
  • a turbomachinery blade of the type including a root portion, a blade platform, a hollow body portion defining an internal cavity, at least one thin-walled, hollow insert adapted to be positioned within said hollow body portion, and means for delivering a coolant to the interior of said insert, the improvement comprising:
  • said securing means including a bushing adapted to be secured to said insert, and pin means adapted to be secured to both said bushing and to said blade.
  • said insert comprises an impingement insert having an inlet at one end thereof and a multiplicity of relatively small holes therethrough which are directed toward surfaces defining said cavity and are adapted to impinge cooling air thereagainst.
  • said bushing includes a second flange of approximately the same thickness as that of said insert wall, said wall includes a second opening for receiving said second flange and said first and second flanges are adapted to be secured to said insert when said flanges are positioned within said openings.
  • said insert includes a base end which defines said inlet, and said base end is sized so as to provide a slight gap between side walls of said internal cavity and said base end, and said gap is filled with a sealant.
  • said bushing comprises a sleeve adapted to be positioned within said inlet, and said sleeve is adapted to be secured to the inner walls of said insert.
  • a turbomachinery blade comprising a root portion, an airfoil-shaped, hollow body portion defining an internal cavity therein, a blade platform separating said root portion and said hollow body portion and defining the inner bounds of said airfoil, at least one thin-walled, hollow insert positioned within said hollow body portion, passage means adapted to deliver a coolant to the interior of said insert, and means for mechanically securing said insert in a position within said internal cavity, said securing means including a bushing adapted to be secured to said insert and pin means adapted to be secured to both said bushing and to said blade.
  • said insert comprises an impingement insert having an inlet at one end thereof and a multiplicity of relatively small holes therethrough which are directed toward surfaces defining said cavity and are adapted to impinge cooling air thereagainst.
  • said bushing includes at least one flange of approximately the same thickness as that of said insert wall, said wall includes an opening for receiving said flange, and said blade further includes means for securing said flange in said opening.
  • said bushing includes a second flange of approximately the same thickness as that of said insert wall, said wall includes a second opening for receiving said second flange, and means for securing said second flange in said second opening.
  • turbomachinery blade recited in claim 11 wherein said spool includes at least one flat side.
  • turbomachinery blade recited in claim 8 further including means for dividing said internal cavity into at least two separate cavities, each of said cavities is provided with one of said impingement inserts, and each of said cavities includes means for mechanically securing said insert to said blade.
  • said root portion of said blade includes at least one hole
  • said bushing includes at least one holeadapted to align with said hole in said root portion
  • said pin means comprise a pin adapted to be positioned within each of said holes.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US346422A 1973-03-30 1973-03-30 Turbomachinery blade cooling insert retainers Expired - Lifetime US3867068A (en)

Priority Applications (10)

Application Number Priority Date Filing Date Title
US346422A US3867068A (en) 1973-03-30 1973-03-30 Turbomachinery blade cooling insert retainers
GB1120974A GB1467374A (en) 1973-03-30 1974-03-13 Hollow turbomachinery blades
CA195,247A CA994672A (en) 1973-03-30 1974-03-18 Turbomachinery blade cooling insert retainers
DE2413292A DE2413292A1 (de) 1973-03-30 1974-03-20 Blattkuehleinsatzhalter fuer turbomaschinen
IT49539/74A IT1003841B (it) 1973-03-30 1974-03-22 Elementi di ritegno per inserti di raffreddamento per palette di turbomacchine
SE7404215A SE390433B (sv) 1973-03-30 1974-03-28 Anordning vid for gasturbiner avsett rotorblad
NL7404257A NL7404257A (xx) 1973-03-30 1974-03-28
JP49034775A JPS5026103A (xx) 1973-03-30 1974-03-29
BE142677A BE813090A (fr) 1973-03-30 1974-03-29 Dispositif de fixation de pieces rapportees d'aubes de turbo-machines
FR7411191A FR2223550B1 (xx) 1973-03-30 1974-03-29

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US346422A US3867068A (en) 1973-03-30 1973-03-30 Turbomachinery blade cooling insert retainers

Publications (1)

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US3867068A true US3867068A (en) 1975-02-18

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Family Applications (1)

Application Number Title Priority Date Filing Date
US346422A Expired - Lifetime US3867068A (en) 1973-03-30 1973-03-30 Turbomachinery blade cooling insert retainers

Country Status (10)

Country Link
US (1) US3867068A (xx)
JP (1) JPS5026103A (xx)
BE (1) BE813090A (xx)
CA (1) CA994672A (xx)
DE (1) DE2413292A1 (xx)
FR (1) FR2223550B1 (xx)
GB (1) GB1467374A (xx)
IT (1) IT1003841B (xx)
NL (1) NL7404257A (xx)
SE (1) SE390433B (xx)

Cited By (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
US4019831A (en) * 1974-09-05 1977-04-26 Brown Boveri Sulzer Turbomachinery Ltd. Cooled rotor blade for a gas turbine
US4177010A (en) * 1977-01-04 1979-12-04 Rolls-Royce Limited Cooled rotor blade for a gas turbine engine
US4321010A (en) * 1978-08-17 1982-03-23 Rolls-Royce Limited Aerofoil member for a gas turbine engine
US4347037A (en) * 1979-02-05 1982-08-31 The Garrett Corporation Laminated airfoil and method for turbomachinery
DE3110096A1 (de) * 1981-03-16 1982-09-23 MTU Motoren- und Turbinen-Union München GmbH, 8000 München "turbinenschaufel, insbesondere turbinenlaufschaufel fuer gasturbinentriebwerke"
US4484859A (en) * 1980-01-17 1984-11-27 Rolls-Royce Limited Rotor blade for a gas turbine engine
US4589824A (en) * 1977-10-21 1986-05-20 United Technologies Corporation Rotor blade having a tip cap end closure
US5125798A (en) * 1990-04-13 1992-06-30 General Electric Company Method and apparatus for cooling air flow at gas turbine bucket trailing edge tip
US5183385A (en) * 1990-11-19 1993-02-02 General Electric Company Turbine blade squealer tip having air cooling holes contiguous with tip interior wall surface
US5207556A (en) * 1992-04-27 1993-05-04 General Electric Company Airfoil having multi-passage baffle
EP1006263A1 (de) * 1998-11-30 2000-06-07 Asea Brown Boveri AG Schaufelkühlung
US6193465B1 (en) 1998-09-28 2001-02-27 General Electric Company Trapped insert turbine airfoil
US6257828B1 (en) * 1997-07-29 2001-07-10 Siemens Aktiengesellschaft Turbine blade and method of producing a turbine blade
US6450759B1 (en) * 2001-02-16 2002-09-17 General Electric Company Gas turbine nozzle vane insert and methods of installation
US20030194320A1 (en) * 2002-02-19 2003-10-16 The Boeing Company Method of fabricating a shape memory alloy damped structure
US6752594B2 (en) 2002-02-07 2004-06-22 The Boeing Company Split blade frictional damper
GB2405186A (en) * 2003-08-20 2005-02-23 Rolls Royce Plc A hollow turbine blade with internal damping element
EP1647672A2 (en) * 2004-10-18 2006-04-19 United Technologies Corporation Airfoil with impingement cooling of a large fillet
US20090081048A1 (en) * 2006-04-21 2009-03-26 Beeck Alexander R Turbine Blade for a Turbine
US8052391B1 (en) * 2009-03-25 2011-11-08 Florida Turbine Technologies, Inc. High temperature turbine rotor blade
US20120134845A1 (en) * 2010-11-29 2012-05-31 Alexander Anatolievich Khanin Blade for a gas turbine, method for manufacturing a turbine blade, and gas turbine with a blade
US8336206B1 (en) * 2009-03-16 2012-12-25 Florida Turbine Technologies, Inc. Process of forming a high temperature turbine rotor blade
WO2013130575A1 (en) * 2012-02-29 2013-09-06 Solar Turbines Incorporated Turbine nozzle insert
WO2015193068A1 (de) * 2014-06-17 2015-12-23 Siemens Aktiengesellschaft Turbinenschaufel und turbine
US20160017724A1 (en) * 2013-04-03 2016-01-21 United Technologies Corporation Variable thickness trailing edge cavity and method of making
WO2016058900A1 (de) * 2014-10-14 2016-04-21 Siemens Aktiengesellschaft Turbinenschaufel mit innenmodul und verfahren zur herstellung einer turbinenschaufel
US20160208621A1 (en) * 2013-09-06 2016-07-21 United Technologies Corporation Gas turbine engine airfoil with wishbone baffle cooling scheme
US10400616B2 (en) 2013-07-19 2019-09-03 General Electric Company Turbine nozzle with impingement baffle
US10626740B2 (en) 2016-12-08 2020-04-21 General Electric Company Airfoil trailing edge segment
US11220916B2 (en) 2020-01-22 2022-01-11 General Electric Company Turbine rotor blade with platform with non-linear cooling passages by additive manufacture
US11242760B2 (en) * 2020-01-22 2022-02-08 General Electric Company Turbine rotor blade with integral impingement sleeve by additive manufacture
US11248471B2 (en) 2020-01-22 2022-02-15 General Electric Company Turbine rotor blade with angel wing with coolant transfer passage between adjacent wheel space portions by additive manufacture
US11492908B2 (en) 2020-01-22 2022-11-08 General Electric Company Turbine rotor blade root with hollow mount with lattice support structure by additive manufacture

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2476207A1 (fr) * 1980-02-19 1981-08-21 Snecma Perfectionnement aux aubes de turbines refroidies
JPS5798387U (xx) * 1980-12-10 1982-06-17
GB2121483B (en) * 1982-06-08 1985-02-13 Rolls Royce Cooled turbine blade for a gas turbine engine

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2779565A (en) * 1948-01-05 1957-01-29 Bruno W Bruckmann Air cooling of turbine blades
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2779565A (en) * 1948-01-05 1957-01-29 Bruno W Bruckmann Air cooling of turbine blades
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure

Cited By (50)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4019831A (en) * 1974-09-05 1977-04-26 Brown Boveri Sulzer Turbomachinery Ltd. Cooled rotor blade for a gas turbine
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
US4177010A (en) * 1977-01-04 1979-12-04 Rolls-Royce Limited Cooled rotor blade for a gas turbine engine
US4589824A (en) * 1977-10-21 1986-05-20 United Technologies Corporation Rotor blade having a tip cap end closure
US4321010A (en) * 1978-08-17 1982-03-23 Rolls-Royce Limited Aerofoil member for a gas turbine engine
US4421153A (en) * 1978-08-17 1983-12-20 Rolls-Royce Limited Method of making an aerofoil member for a gas turbine engine
US4347037A (en) * 1979-02-05 1982-08-31 The Garrett Corporation Laminated airfoil and method for turbomachinery
US4484859A (en) * 1980-01-17 1984-11-27 Rolls-Royce Limited Rotor blade for a gas turbine engine
DE3110096A1 (de) * 1981-03-16 1982-09-23 MTU Motoren- und Turbinen-Union München GmbH, 8000 München "turbinenschaufel, insbesondere turbinenlaufschaufel fuer gasturbinentriebwerke"
US5125798A (en) * 1990-04-13 1992-06-30 General Electric Company Method and apparatus for cooling air flow at gas turbine bucket trailing edge tip
US5183385A (en) * 1990-11-19 1993-02-02 General Electric Company Turbine blade squealer tip having air cooling holes contiguous with tip interior wall surface
US5207556A (en) * 1992-04-27 1993-05-04 General Electric Company Airfoil having multi-passage baffle
US6257828B1 (en) * 1997-07-29 2001-07-10 Siemens Aktiengesellschaft Turbine blade and method of producing a turbine blade
US6193465B1 (en) 1998-09-28 2001-02-27 General Electric Company Trapped insert turbine airfoil
EP1006263A1 (de) * 1998-11-30 2000-06-07 Asea Brown Boveri AG Schaufelkühlung
US6328532B1 (en) 1998-11-30 2001-12-11 Alstom Blade cooling
US6450759B1 (en) * 2001-02-16 2002-09-17 General Electric Company Gas turbine nozzle vane insert and methods of installation
US6752594B2 (en) 2002-02-07 2004-06-22 The Boeing Company Split blade frictional damper
US20030194320A1 (en) * 2002-02-19 2003-10-16 The Boeing Company Method of fabricating a shape memory alloy damped structure
US6699015B2 (en) 2002-02-19 2004-03-02 The Boeing Company Blades having coolant channels lined with a shape memory alloy and an associated fabrication method
US6886622B2 (en) 2002-02-19 2005-05-03 The Boeing Company Method of fabricating a shape memory alloy damped structure
GB2405186B (en) * 2003-08-20 2005-10-26 Rolls Royce Plc A component with internal damping
US7070390B2 (en) 2003-08-20 2006-07-04 Rolls-Royce Plc Component with internal damping
GB2405186A (en) * 2003-08-20 2005-02-23 Rolls Royce Plc A hollow turbine blade with internal damping element
US20050047918A1 (en) * 2003-08-20 2005-03-03 Rolls-Royce Plc Component with internal damping
US7220103B2 (en) 2004-10-18 2007-05-22 United Technologies Corporation Impingement cooling of large fillet of an airfoil
US20060083613A1 (en) * 2004-10-18 2006-04-20 United Technologies Corporation Impingement cooling of large fillet of an airfoil
EP1647672A3 (en) * 2004-10-18 2006-09-06 United Technologies Corporation Airfoil with impingement cooling of a large fillet
EP1647672A2 (en) * 2004-10-18 2006-04-19 United Technologies Corporation Airfoil with impingement cooling of a large fillet
US20090081048A1 (en) * 2006-04-21 2009-03-26 Beeck Alexander R Turbine Blade for a Turbine
US8047001B2 (en) * 2006-04-21 2011-11-01 Siemens Aktiengesellschaft Media mixing insert for turbine blade in turbine engine
US8336206B1 (en) * 2009-03-16 2012-12-25 Florida Turbine Technologies, Inc. Process of forming a high temperature turbine rotor blade
US8052391B1 (en) * 2009-03-25 2011-11-08 Florida Turbine Technologies, Inc. High temperature turbine rotor blade
US9188011B2 (en) * 2010-11-29 2015-11-17 Alstom Technology Ltd. Blade for a gas turbine, method for manufacturing a turbine blade, and gas turbine with a blade
US20120134845A1 (en) * 2010-11-29 2012-05-31 Alexander Anatolievich Khanin Blade for a gas turbine, method for manufacturing a turbine blade, and gas turbine with a blade
WO2013130575A1 (en) * 2012-02-29 2013-09-06 Solar Turbines Incorporated Turbine nozzle insert
US20160017724A1 (en) * 2013-04-03 2016-01-21 United Technologies Corporation Variable thickness trailing edge cavity and method of making
US10400616B2 (en) 2013-07-19 2019-09-03 General Electric Company Turbine nozzle with impingement baffle
US10975705B2 (en) 2013-09-06 2021-04-13 Raytheon Technologies Corporation Gas turbine engine airfoil with wishbone baffle cooling scheme
US20160208621A1 (en) * 2013-09-06 2016-07-21 United Technologies Corporation Gas turbine engine airfoil with wishbone baffle cooling scheme
US10487668B2 (en) * 2013-09-06 2019-11-26 United Technologies Corporation Gas turbine engine airfoil with wishbone baffle cooling scheme
EP2957724A1 (de) * 2014-06-17 2015-12-23 Siemens Aktiengesellschaft Turbinenschaufel und Turbine
WO2015193068A1 (de) * 2014-06-17 2015-12-23 Siemens Aktiengesellschaft Turbinenschaufel und turbine
WO2016058900A1 (de) * 2014-10-14 2016-04-21 Siemens Aktiengesellschaft Turbinenschaufel mit innenmodul und verfahren zur herstellung einer turbinenschaufel
US10626740B2 (en) 2016-12-08 2020-04-21 General Electric Company Airfoil trailing edge segment
US11220916B2 (en) 2020-01-22 2022-01-11 General Electric Company Turbine rotor blade with platform with non-linear cooling passages by additive manufacture
US11242760B2 (en) * 2020-01-22 2022-02-08 General Electric Company Turbine rotor blade with integral impingement sleeve by additive manufacture
US11248471B2 (en) 2020-01-22 2022-02-15 General Electric Company Turbine rotor blade with angel wing with coolant transfer passage between adjacent wheel space portions by additive manufacture
US11492908B2 (en) 2020-01-22 2022-11-08 General Electric Company Turbine rotor blade root with hollow mount with lattice support structure by additive manufacture

Also Published As

Publication number Publication date
DE2413292A1 (de) 1974-10-10
FR2223550A1 (xx) 1974-10-25
SE390433B (sv) 1976-12-20
CA994672A (en) 1976-08-10
BE813090A (fr) 1974-09-30
IT1003841B (it) 1976-06-10
GB1467374A (en) 1977-03-16
FR2223550B1 (xx) 1978-01-13
JPS5026103A (xx) 1975-03-19
NL7404257A (xx) 1974-10-02

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