US3182955A - Construction of turbomachinery blade elements - Google Patents

Construction of turbomachinery blade elements Download PDF

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Publication number
US3182955A
US3182955A US144740A US14474061A US3182955A US 3182955 A US3182955 A US 3182955A US 144740 A US144740 A US 144740A US 14474061 A US14474061 A US 14474061A US 3182955 A US3182955 A US 3182955A
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Prior art keywords
blade
lugs
plate
segments
stage
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Expired - Lifetime
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US144740A
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John A C Hyde
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Ruston and Hornsby Ltd
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Ruston and Hornsby Ltd
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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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • 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
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F05D2240/10—Stators
    • 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
    • F05D2240/81—Cooled platforms
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/4932—Turbomachine making
    • Y10T29/49321—Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member

Definitions

  • This invention relates to axial flow turbomachinery such as turbines and compressors having a working fluid passing through annular rows or stages of rotor and stator blades or vanes, particularly gas turbines operating at high temperatures when thermal expansion and distortion may adversely affect working clearances and alignment of such blades or vanes and their associated structure.
  • the invention more particularly relates to the improved construction and connecting and reconditioning of annular stages of rotor and stator blades and their associated structure.
  • blades in turbomachinery by manufacturing units consisting of at least one blade and at least one associated plate generally known in the art as a platform or a shroud according to its locationat either the root end or the tip of a rotor or stator blade.
  • platforms and shrouds the said platforms and shrouds will be hereinafter referred to as plates.
  • Each unit consisting of at least one blade and at least one associated plate will be hereinafter referred to as a segment.
  • annular stage is built up of a plurality of segments with the substantially axial edges of each plate abutting the corresponding edges'of its neighbouring plates so that when the annular stage is complete the said plates form a continuous ring, ir if each segment has two plates, one at each end of the blade, the said plates form two continuous concentric rings;
  • the manner of locating and connecting the said plates has been given much thought in view of the necessity of obtaining and maintaining accurate blade alignment against thermal expansion and contraction.
  • the invention consists in a plurality of segments as defined above and connected to form an annular row or stage by means of pairs of lugs spaced apart radially with respect to the stage by grooves formed on the substantially Y axial edges of each plate which abuts the corresponding edges of its neighbouring plates within the stage, the inner' lugs with respect to the flow of working fluid being in contiguous relationship with all the inner lugs in a com tinuous ring, and the outer lugs with respect to the flow of working fiuid being butt-welded to the neighbouring outer lugs.
  • the invention further consists in a plurality of segments constructed and connected as described wherein each pair of abutting lugs define between-them a channel for a flow of fluid coolant, which may also be conveyed to the interior of the blades by means of passages formed in the blades and communicating ducts formed in the plates between channel and blades.
  • the invention also consists in an annular stage ofseg ments constructed as described wherein the inner logs with respect to the flow of working fluid are located in contiguous relationship with each otherby means of cooperating flanges 'formed in theirabutting faces, the said flanges being secured against movement and distortion relative to each other by conventional securing means, for
  • the plates may be provided with brackets, flanges, or other fixing means whereby the segments may be located and secured within the turbomachinery.
  • FIGURE 1 is an isometric view of two. segments conj nected according to the invention, :and
  • FIGURE 2 is a partially sectioned detail of alternative constructions.
  • each segment consisting of a blade 11 and two plates 12 and 13 defining between them a passage ;for the flow of working fluid.
  • the lugs 14, 15, are spaced apart radially with respectto the stage by a groove 16.
  • the outer lugs 15 are butt-welded to their neighbouring outer lugs 15 as at 17, and the inner lugs 14 are in contiguous butting relationship with their neighbouring inner lugs 14.
  • FIGURE 2 also shows an example of alternative construction for inner lugs 14 as a further safeguard against movement or distortion of 'the lugs 14 relative to each other.
  • Flanges formed in their abutting faces may be secured together by a conventional dowel 21.
  • FIGURE 1 also'shows an example of a bracket 22 by which the segment may be located and secured within the turbomachinery casing (not shown).
  • the bracket 22 may be designed to prevent radial or other displacement relative to the rotor stages and the turbomachinery casing.
  • a turbo machinery stator blade stage comprising: a plurality of blade segments each including at least one generally radial'blade and a plate which generally a'segment of an annulus on at least one end of said blade and said plate having a pair of axially opposed edges; and means for connecting said blade segments to form a stator blade stage, said means including inner and outer, radially spaced lugs on the axial edges of said plate, said outer lug being crank-shaped in cross-section and having a portion offset radially from said plate, the contiguous inner lugs of adjacent segments abutting each other and the contiguous outer lugs of adjacent segments being welded wherein said plates and said blades have communicating ducts formed therein and wherein said inner and outer lugs define between them a channel for the flow of fluid coolant conveyed to the interior of said blades through said communicating ducts.

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

Description

y 1965 J. A. c. HYDE 3,182,955
CONSTRUCTION OF TURBOMACHINERY BLADE ELEMENTS Filed 001;. 12. 1961 INVENTOR fiale... BY m Y ATTORNEYS United States PatentO 3,182,955 CONSTRUCTION OF TURBOMACHINERY BLADE ELEMENTS John A. C. Hyde, Lincoln, England, assignor to Boston & Hornsby Limited, Lincoln, England, a corporation of Great Britain Filed Oct." 12, 1961, Ser. No. 144,740 7 Claims priority, application Great Britain, Get. 29, "1960, 37,254/ 60 3 Claims. (Cl. 253-391) This invention relates to axial flow turbomachinery such as turbines and compressors having a working fluid passing through annular rows or stages of rotor and stator blades or vanes, particularly gas turbines operating at high temperatures when thermal expansion and distortion may adversely affect working clearances and alignment of such blades or vanes and their associated structure. The invention more particularly relates to the improved construction and connecting and reconditioning of annular stages of rotor and stator blades and their associated structure.
-It isknown to construct annular stages of rotor and stator blades or vanes, hereinafter referred to as blades, in turbomachinery by manufacturing units consisting of at least one blade and at least one associated plate generally known in the art as a platform or a shroud according to its locationat either the root end or the tip of a rotor or stator blade. As the invention applies equally to platforms and shrouds, the said platforms and shrouds will be hereinafter referred to as plates. Each unit consisting of at least one blade and at least one associated plate will be hereinafter referred to as a segment. In the known art, an annular stage is built up of a plurality of segments with the substantially axial edges of each plate abutting the corresponding edges'of its neighbouring plates so that when the annular stage is complete the said plates form a continuous ring, ir if each segment has two plates, one at each end of the blade, the said plates form two continuous concentric rings; The manner of locating and connecting the said plates has been given much thought in view of the necessity of obtaining and maintaining accurate blade alignment against thermal expansion and contraction.
The invention consists in a plurality of segments as defined above and connected to form an annular row or stage by means of pairs of lugs spaced apart radially with respect to the stage by grooves formed on the substantially Y axial edges of each plate which abuts the corresponding edges of its neighbouring plates within the stage, the inner' lugs with respect to the flow of working fluid being in contiguous relationship with all the inner lugs in a com tinuous ring, and the outer lugs with respect to the flow of working fiuid being butt-welded to the neighbouring outer lugs.
The invention further consists in a plurality of segments constructed and connected as described wherein each pair of abutting lugs define between-them a channel for a flow of fluid coolant, which may also be conveyed to the interior of the blades by means of passages formed in the blades and communicating ducts formed in the plates between channel and blades.
The invention also consists in an annular stage ofseg ments constructed as described wherein the inner logs with respect to the flow of working fluid are located in contiguous relationship with each otherby means of cooperating flanges 'formed in theirabutting faces, the said flanges being secured against movement and distortion relative to each other by conventional securing means, for
' example, studs or dowels.
The plates may be provided with brackets, flanges, or other fixing means whereby the segments may be located and secured within the turbomachinery.
ice
The following description relates to the drawings accompanying the specification, which show by way of example the application of the invention to the construction of an annular stage of stator segments, FIGURE 1 is an isometric view of two. segments conj nected according to the invention, :and
FIGURE 2 is a partially sectioned detail of alternative constructions.
Referring to FIGURE 1, two segments are shown connected, each segment consisting of a blade 11 and two plates 12 and 13 defining between them a passage ;for the flow of working fluid. On the substantially axial edges of each plate 12, 13, which abuts the corresponding edges of neighbouring plates Within the stage are formed pairs of lugs; these are termed, with respect to their nearness or remoteness from the working fluid, inner lug 14 and outer lug 15. The lugs 14, 15, are spaced apart radially with respectto the stage by a groove 16. The outer lugs 15 are butt-welded to their neighbouring outer lugs 15 as at 17, and the inner lugs 14 are in contiguous butting relationship with their neighbouring inner lugs 14.
The advantages of this construction are: It enables a and contraction and the relative positions of the blades within the annular stage will be maintained within very fine limits. Another advantage is that if in the course of service one of the blades becomes unserviceable, it may be cut out and a new blade welded or otherwise secured into the segment, again without disturbing blade alignment, whereas with conventional segment assembly the whole segment would be rejected. A further advantage is that with the meeting of two lugs 14 and two lugs 15, the two grooves 16 form a channel 18 (see FIGURE 2) which may be used for a flow of coolant. If desired, ducts 19 may be formed in plates 12 to form communicating ducts, to convey coolant from channel 18 to ducts 29 formed in the interior of blades 11.
FIGURE 2 also shows an example of alternative construction for inner lugs 14 as a further safeguard against movement or distortion of 'the lugs 14 relative to each other. Flanges formed in their abutting faces may be secured together by a conventional dowel 21.
" FIGURE 1 also'shows an example of a bracket 22 by which the segment may be located and secured within the turbomachinery casing (not shown). The bracket 22 may be designed to prevent radial or other displacement relative to the rotor stages and the turbomachinery casing.
What I claim is:
l. A turbo machinery stator blade stage comprising: a plurality of blade segments each including at least one generally radial'blade and a plate which generally a'segment of an annulus on at least one end of said blade and said plate having a pair of axially opposed edges; and means for connecting said blade segments to form a stator blade stage, said means including inner and outer, radially spaced lugs on the axial edges of said plate, said outer lug being crank-shaped in cross-section and having a portion offset radially from said plate, the contiguous inner lugs of adjacent segments abutting each other and the contiguous outer lugs of adjacent segments being welded wherein said plates and said blades have communicating ducts formed therein and wherein said inner and outer lugs define between them a channel for the flow of fluid coolant conveyed to the interior of said blades through said communicating ducts. 5 3. The turbomachinery blade stage recited in claim 2, wherein said inner lugs include cooperating flanges formed in their abuttting faces for locating them in contiguous relationship with each other and means for securing said flanges against movement and distortion relative to each 10 other.
References Cited by the Examiner UNITED STATES PATENTS Hcppner 253-77 Bloomberg 253-77 Bodger 253-77 Sollinger 253-77 Greenwald 29-1568 Galliot 230-122 Wayne 25.3-39.15 Morley 253-78 Cutler 25.3-39.1 Klompas et a1 253-77 Hockert et a1. 253-77 Stark 253-3915 JOSEPH H. BRANSON, JR., Primary Examiner. 1,470,508 10 23 Steerstrup 253.47 15 WALTER BERLOWITZ, Emmi/wr- UNITED STATES PATENT OFFICE CERTIFICATE OF CORRECTION Patent No. 3,182,955 May 11, 1965 John A. C. Hyde It is hereby certified that error appears in the above numbered patent requiring correction and that the said Letters Patent should read as corrected below.
Column 1, line 36, for "ir" read or line 53, for "fiuid" read fluid column 2, line 24, for "criticial" read critical line 61, after "which" insert is Signed and sealed this 28th day of December 1965.
(SEAL) Attest:
ERNEST W. SWIDER EDWARD J. BRENNER Attesting Officer Commissioner of Patents

Claims (1)

1. A TURBO MACHINERY STATOR BLADE STAGE COMPRISING: A PLURALITY OF BLADE SEGMENTS EACH INCLUDING AT LEAST ONE GENERALLY RADIAL BLADE AND PLATE WHICH GENERALLY A SEGMENT OF AN ANNULUS ON AT LEAST ONE END OF SAID BLADE AND SAID PLATE HAVING A PAIR OF AXIALLY OPPOSED EDGES; AND MEANS FOR CONNECTING SAID BLADE SEGMENTS TO FORM A STATOR BLADE STAGE, SAID MEANS INCLUDING INNER AND OUTER, RADIALLY SPACED LUGS ON THE AXIAL EDGES OF SAID PLATE, SAID OUTER LUG BEING CRANK-SHAPED IN CROSS-SECTION AND HAVING A PORTION OFFSET RADIALLY FROM SAID PLATE, THE CONTIGUOUS INNER LUGS OF ADJACENT SEGMENTS ABUTTING EACH OTHER AND THE CONTIGUOUS OUTER LUGS OF ADJACENT SEGMENTS BEING WELDED TO EACH OTHER.
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Cited By (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3365173A (en) * 1966-02-28 1968-01-23 Gen Electric Stator structure
US3527544A (en) * 1968-12-12 1970-09-08 Gen Motors Corp Cooled blade shroud
US3628226A (en) * 1970-03-16 1971-12-21 Aerojet General Co Method of making hollow compressor blades
US4015910A (en) * 1976-03-09 1977-04-05 The United States Of America As Represented By The Secretary Of The Air Force Bolted paired vanes for turbine
US4688988A (en) * 1984-12-17 1987-08-25 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US4812107A (en) * 1985-02-28 1989-03-14 Bbc Brown, Boveri & Company, Ltd. Method of manufacturing a control wheel for the high-pressure rotor of a steam turbine
DE2609702C1 (en) * 1975-03-14 1989-05-03 Rolls Royce Jet guide vane for a gas turbine engine
US4850090A (en) * 1987-07-22 1989-07-25 Rolls-Royce Plc Method of manufacture of an axial flow compressor stator assembly
EP0357984A1 (en) * 1988-08-31 1990-03-14 Westinghouse Electric Corporation Gas turbine with film cooling of turbine vane shrouds
FR2637320A1 (en) * 1988-09-30 1990-04-06 Rolls Royce Plc PROFILED BLADE OF TURBINE
EP0921273A1 (en) * 1997-06-11 1999-06-09 Mitsubishi Heavy Industries, Ltd. Rotor for gas turbines
EP1022437A1 (en) * 1999-01-19 2000-07-26 Siemens Aktiengesellschaft Construction element for use in a thermal machine
US6210111B1 (en) * 1998-12-21 2001-04-03 United Technologies Corporation Turbine blade with platform cooling
US6553665B2 (en) * 2000-03-08 2003-04-29 General Electric Company Stator vane assembly for a turbine and method for forming the assembly
EP1760268A2 (en) * 2005-08-30 2007-03-07 General Electric Company Apparatus for controlling contact within stator assemblies
EP1892383A1 (en) * 2006-08-24 2008-02-27 Siemens Aktiengesellschaft Gas turbine blade with cooled platform
EP1995409A2 (en) 2007-05-22 2008-11-26 United Technologies Corporation Repair method for turbine vanes
US20090067987A1 (en) * 2007-08-06 2009-03-12 United Technologies Corporation Airfoil replacement repair
US20090274562A1 (en) * 2008-05-02 2009-11-05 United Technologies Corporation Coated turbine-stage nozzle segments
US20100180417A1 (en) * 2009-01-20 2010-07-22 United Technologies Corporation Replacement of part of engine case with dissimilar material
JP2010242750A (en) * 2009-03-31 2010-10-28 General Electric Co <Ge> Feeding film cooling hole from seal slot
US20120121384A1 (en) * 2010-11-15 2012-05-17 Mtu Aero Engines Gmbh Rotor and method for manufacturing a rotor for a turbo machine
US20130052020A1 (en) * 2011-08-23 2013-02-28 General Electric Company Coupled blade platforms and methods of sealing
US8763403B2 (en) 2010-11-19 2014-07-01 United Technologies Corporation Method for use with annular gas turbine engine component
US20140271171A1 (en) * 2013-03-15 2014-09-18 Edward Len Miller Compressor airfoil
US20160153297A1 (en) * 2013-07-30 2016-06-02 United Tehnologies Corporation Gas turbine engine turbine vane ring arrangement
US9631517B2 (en) 2012-12-29 2017-04-25 United Technologies Corporation Multi-piece fairing for monolithic turbine exhaust case
US20190085706A1 (en) * 2017-09-18 2019-03-21 General Electric Company Turbine engine airfoil assembly
EP4202186A1 (en) * 2021-12-27 2023-06-28 Rolls-Royce plc Turbine blade

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US5176496A (en) * 1991-09-27 1993-01-05 General Electric Company Mounting arrangements for turbine nozzles

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1470508A (en) * 1921-12-23 1923-10-09 Gen Electric Method of manufacturing turbine elements
US2398140A (en) * 1943-12-08 1946-04-09 Armstrong Siddeley Motors Ltd Bladed rotor
US2500745A (en) * 1944-09-21 1950-03-14 Gen Electric Bucket structure for high-temperature turbomachines
US2510734A (en) * 1946-04-06 1950-06-06 United Aircraft Corp Turbine or compressor rotor
US2603453A (en) * 1946-09-11 1952-07-15 Curtiss Wright Corp Cooling means for turbines
US2639119A (en) * 1947-11-14 1953-05-19 Lockheed Aircraft Corp Rotor blade attachment means and method
US2646209A (en) * 1948-05-21 1953-07-21 Galliot Jules Andre Norbert Turbine driven multistage compressor
US2819870A (en) * 1955-04-18 1958-01-14 Oleh A Wayne Sheet metal blade base
US2833463A (en) * 1953-11-06 1958-05-06 Rolls Royce Stator construction for axial flow compressor
US2924425A (en) * 1953-02-02 1960-02-09 Bristol Aero Engines Ltd Aerofoil-section bladed structures
US2931622A (en) * 1956-12-24 1960-04-05 Orenda Engines Ltd Rotor construction for gas turbine engines
US2945673A (en) * 1951-10-31 1960-07-19 Gen Motors Corp Segmented stator ring assembly
US3044745A (en) * 1956-11-20 1962-07-17 Rolls Royce Turbine and compressor blades

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1470508A (en) * 1921-12-23 1923-10-09 Gen Electric Method of manufacturing turbine elements
US2398140A (en) * 1943-12-08 1946-04-09 Armstrong Siddeley Motors Ltd Bladed rotor
US2500745A (en) * 1944-09-21 1950-03-14 Gen Electric Bucket structure for high-temperature turbomachines
US2510734A (en) * 1946-04-06 1950-06-06 United Aircraft Corp Turbine or compressor rotor
US2603453A (en) * 1946-09-11 1952-07-15 Curtiss Wright Corp Cooling means for turbines
US2639119A (en) * 1947-11-14 1953-05-19 Lockheed Aircraft Corp Rotor blade attachment means and method
US2646209A (en) * 1948-05-21 1953-07-21 Galliot Jules Andre Norbert Turbine driven multistage compressor
US2945673A (en) * 1951-10-31 1960-07-19 Gen Motors Corp Segmented stator ring assembly
US2924425A (en) * 1953-02-02 1960-02-09 Bristol Aero Engines Ltd Aerofoil-section bladed structures
US2833463A (en) * 1953-11-06 1958-05-06 Rolls Royce Stator construction for axial flow compressor
US2819870A (en) * 1955-04-18 1958-01-14 Oleh A Wayne Sheet metal blade base
US3044745A (en) * 1956-11-20 1962-07-17 Rolls Royce Turbine and compressor blades
US2931622A (en) * 1956-12-24 1960-04-05 Orenda Engines Ltd Rotor construction for gas turbine engines

Cited By (43)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3365173A (en) * 1966-02-28 1968-01-23 Gen Electric Stator structure
US3527544A (en) * 1968-12-12 1970-09-08 Gen Motors Corp Cooled blade shroud
US3628226A (en) * 1970-03-16 1971-12-21 Aerojet General Co Method of making hollow compressor blades
DE2609702C1 (en) * 1975-03-14 1989-05-03 Rolls Royce Jet guide vane for a gas turbine engine
US4015910A (en) * 1976-03-09 1977-04-05 The United States Of America As Represented By The Secretary Of The Air Force Bolted paired vanes for turbine
US4688988A (en) * 1984-12-17 1987-08-25 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US4812107A (en) * 1985-02-28 1989-03-14 Bbc Brown, Boveri & Company, Ltd. Method of manufacturing a control wheel for the high-pressure rotor of a steam turbine
US4850090A (en) * 1987-07-22 1989-07-25 Rolls-Royce Plc Method of manufacture of an axial flow compressor stator assembly
EP0357984A1 (en) * 1988-08-31 1990-03-14 Westinghouse Electric Corporation Gas turbine with film cooling of turbine vane shrouds
US4948338A (en) * 1988-09-30 1990-08-14 Rolls-Royce Plc Turbine blade with cooled shroud abutment surface
FR2637320A1 (en) * 1988-09-30 1990-04-06 Rolls Royce Plc PROFILED BLADE OF TURBINE
EP0921273A1 (en) * 1997-06-11 1999-06-09 Mitsubishi Heavy Industries, Ltd. Rotor for gas turbines
EP0921273A4 (en) * 1997-06-11 2001-01-24 Mitsubishi Heavy Ind Ltd Rotor for gas turbines
US6210111B1 (en) * 1998-12-21 2001-04-03 United Technologies Corporation Turbine blade with platform cooling
EP1022437A1 (en) * 1999-01-19 2000-07-26 Siemens Aktiengesellschaft Construction element for use in a thermal machine
US6553665B2 (en) * 2000-03-08 2003-04-29 General Electric Company Stator vane assembly for a turbine and method for forming the assembly
EP1760268A2 (en) * 2005-08-30 2007-03-07 General Electric Company Apparatus for controlling contact within stator assemblies
EP1760268A3 (en) * 2005-08-30 2011-12-21 General Electric Company Apparatus for controlling contact within stator assemblies
EP1892383A1 (en) * 2006-08-24 2008-02-27 Siemens Aktiengesellschaft Gas turbine blade with cooled platform
WO2008022830A1 (en) * 2006-08-24 2008-02-28 Siemens Aktiengesellschaft Gas turbine blade having a cooled platform
US20080289179A1 (en) * 2007-05-22 2008-11-27 United Technologies Corporation Split vane repair
EP1995409A2 (en) 2007-05-22 2008-11-26 United Technologies Corporation Repair method for turbine vanes
US8220150B2 (en) 2007-05-22 2012-07-17 United Technologies Corporation Split vane cluster repair method
US20090067987A1 (en) * 2007-08-06 2009-03-12 United Technologies Corporation Airfoil replacement repair
US7798773B2 (en) 2007-08-06 2010-09-21 United Technologies Corporation Airfoil replacement repair
US20090274562A1 (en) * 2008-05-02 2009-11-05 United Technologies Corporation Coated turbine-stage nozzle segments
US20100180417A1 (en) * 2009-01-20 2010-07-22 United Technologies Corporation Replacement of part of engine case with dissimilar material
US8245399B2 (en) 2009-01-20 2012-08-21 United Technologies Corporation Replacement of part of engine case with dissimilar material
JP2010242750A (en) * 2009-03-31 2010-10-28 General Electric Co <Ge> Feeding film cooling hole from seal slot
US20120121384A1 (en) * 2010-11-15 2012-05-17 Mtu Aero Engines Gmbh Rotor and method for manufacturing a rotor for a turbo machine
US8888456B2 (en) * 2010-11-15 2014-11-18 Mtu Aero Engines Gmbh Rotor and method for manufacturing a rotor for a turbo machine
US8763403B2 (en) 2010-11-19 2014-07-01 United Technologies Corporation Method for use with annular gas turbine engine component
US8888459B2 (en) * 2011-08-23 2014-11-18 General Electric Company Coupled blade platforms and methods of sealing
US20130052020A1 (en) * 2011-08-23 2013-02-28 General Electric Company Coupled blade platforms and methods of sealing
US9631517B2 (en) 2012-12-29 2017-04-25 United Technologies Corporation Multi-piece fairing for monolithic turbine exhaust case
US20140271171A1 (en) * 2013-03-15 2014-09-18 Edward Len Miller Compressor airfoil
US9920642B2 (en) * 2013-03-15 2018-03-20 Ansaldo Energia Ip Uk Limited Compressor airfoil
US20160153297A1 (en) * 2013-07-30 2016-06-02 United Tehnologies Corporation Gas turbine engine turbine vane ring arrangement
US10344603B2 (en) * 2013-07-30 2019-07-09 United Technologies Corporation Gas turbine engine turbine vane ring arrangement
US11021980B2 (en) 2013-07-30 2021-06-01 Raytheon Technologies Corporation Gas turbine engine turbine vane ring arrangement
US20190085706A1 (en) * 2017-09-18 2019-03-21 General Electric Company Turbine engine airfoil assembly
EP4202186A1 (en) * 2021-12-27 2023-06-28 Rolls-Royce plc Turbine blade
US11739647B2 (en) 2021-12-27 2023-08-29 Rolls-Royce Plc Turbine blade

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