US3910716A - Gas turbine inlet vane structure utilizing a stable ceramic spherical interface arrangement - Google Patents

Gas turbine inlet vane structure utilizing a stable ceramic spherical interface arrangement Download PDF

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Publication number
US3910716A
US3910716A US472753A US47275374A US3910716A US 3910716 A US3910716 A US 3910716A US 472753 A US472753 A US 472753A US 47275374 A US47275374 A US 47275374A US 3910716 A US3910716 A US 3910716A
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United States
Prior art keywords
end caps
vane
radially
blade
engagement
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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US472753A
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English (en)
Inventor
Jeffrey D Roughgarden
Stephen D Leshnoff
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CBS Corp
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Westinghouse Electric Corp
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Publication date
Application filed by Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Priority to US472753A priority Critical patent/US3910716A/en
Priority to CA225,297A priority patent/CA1001953A/en
Priority to IT23577/75A priority patent/IT1038328B/it
Priority to JP6111575A priority patent/JPS5335205B2/ja
Application granted granted Critical
Publication of US3910716A publication Critical patent/US3910716A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • 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
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/24—Three-dimensional ellipsoidal
    • F05D2250/241—Three-dimensional ellipsoidal spherical
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/21—Oxide ceramics

Definitions

  • ABSTRACT An improved ceramic inlet vane structure for axial flow gas turbines, comprising an array of three ceramic blades arranged in a vane segment. Each blade has a spherical interface between its radially inner and outer ends and their respective supportive end caps. The three points of contact between the inner ends of the three blades and their respective inner end caps in the vane segment generally define a triangle. The points of contact between the outer ends of each the three blades and their respective outer end caps in the vane segment also generally define a triangle.
  • Each group of inner and outer end caps themselves are restrained by a shoe having a single spherical interface with inner and outer shroud members respectively.
  • the four spherical interface points on the inner portions and on the outer portions of each vane segment define a tetrahedron. This configuration provides high stability with a slight freedom of movement within each vane segment and a slight freedom of movement within each blade.
  • An annular arrray of vane segments comprises the inlet nozzle of a gas turbine.
  • This invention relates generally to turbines, and more particularly to ceramic inlet nozzle structures for gas turbines.
  • Gas turbines presently employ integral first stage metal inlet vane segments. As inlet temperatures are increased to increase turbine efficiency, cooling of the metal inlet vanes is necessary. Providing cooling fluid for the metal vanes utilizes some of the power produced by the turbines, hence it decreases the overall efficiency of that unit. Ceramics have been introduced as a high temperature material from which to construct the inlet vane segments. Ceramics, however, are structurally most stable when used in a compressed state. Gas turbine inlet nozzles during operation have an array of forces generated therein that are not strictly compressive. The forces generated therein may be in shear or tension, and are produced because of thermal expansion, movement of adjacent members, and the like.
  • An object of this invention is to overcome the problems associated with the prior art.
  • Another object of this invention is to design a structural arrangement within the inlet nozzle so that those forces generated within the ceramic blades and supportive end caps will be of minimal deleteriousness.
  • Yet another object of this invention is to provide a stable vane segment arrangement that will permit a slight freedom of elongation and rotation for the individual blades within that vane segment.
  • Vane segments an annular array of which comprise the inlet nozzle, themselves are comprised of three radially directed ceramic blades. Each blade has its own supportive radially inner and radially outer ceramic end cap. Each radial end of each of the ceramic blades has a generally hemispherical cavity disposed within it. Each adjacent ceramic end cap has a generally hemispherical cavity aligned with its respective cavity in its radially adjacent blade. A ceramic sphere is disposed in each cavity between each of the ceramic blades and each ceramic end cap. This provides a spherical interface therebetween.
  • the ceramic sphere permits slight freedom of pivotal motion of the blades with respect to one another and with respect to the vane structure itself.
  • Each array of radially inner spheres and each array of radially outer spheres define corners of a triangle.
  • Each array of end caps is restrained by a shoe having a spherical interface with its respective inner and outer shroud.
  • the four inner and four outer points of spherical interface each define a tetrahedron.
  • the radially outer vertex of the outer tetrahedron being compressed radially inwardly, all of the spherical interface points providing a support arrangement that is very stable, and which also permits slight elongation and rotation of its components.
  • FIG. 1 is an exploded perspective view of a portion of an inlet nozzle of a gas turbine, showing a vane segment constructed in accordance with the principles of this invention
  • FIG. 2 is a schematic diagram of a three blade stable vane assembly utilizing a spherical interface arrangement
  • FIG. 3 is a longitudinal sectional view of the middle blade and spherical interface support arrangements of the stable vane assembly
  • FIG. 4 is a longitudinal sectional view of an end blade in a vane segment and a spherical interface arrangement for the blades in a vane segment;
  • FIG. 5 is another embodiment of a spherical interface arrangement for the blades in a vane segment
  • FIG. 6 is yet another embodiment of a spherical interface arrangement for the blades on a vane segment
  • FIG. 6a is still another embodiment of a spherical interface arrangement for the blades on a vane segment.
  • FIG. 7 is a stop arrangement for preventing excessive twist in each blade.
  • FIG. 1 there is shown a portion of an inlet arrangement of an axial flow gas turbine 10 having a stable inlet nozzle arrangement 12.
  • the gas turbine 10 includes a turbine axis 14, an outer cylinder 16, an annular array of vane segments 18 which comprise the inlet nozzle arrangement 12, and at least one rotor disc 20 with an array of rotating blades 22.
  • a fixed plunger biasing arrangement 23 is disposed radially outwardly of each array of vane segments 18 to maintain each array of vane segments 18 in a generally compressed state.
  • the vane segments 18 are preferably constructed from ceramic materials, to withstand the high temperatures, about 2500F, which exist within the inlet nozzle 12 due to the flow of hot gases therethrough from an array of combustion chambers, not shown.
  • the vane segments 18 are maintained in the compressed state because ceramic materials are strongest in that mode.
  • the vane segments 18 are comprised of three individual ceramic airfoil blades 24, each blade 24 having its own supportive radially inner ceramic end caps 26, and their own radially outer ceramic end caps 28.
  • An insulator and supportive member 30 is disposed radially outwardly of the outer ceramic end caps 28 and an insulator and supportive member 32 is disposed radially inwardly of the inner ceramic end caps 26.
  • each ceramic blade has a spherical interface arrangement 25 with its respective adjacent ceramic end cap 26 or 28.
  • the preferred embodiment discloses a rolling spherical interface relationship with each blade 24 and respective end caps 26 and 28, that comprise, with a spherical pivot assembly 27 and 27, radially outwardly and inwardly respectively, of the outer and inner supportive members 30 and 32, a stable vane assembly 12.
  • each ceramic blade 24 has a generally hemispherical cavity 34 disposed therein, as shown in FIGS. 1, 3 and 4.
  • Each outer ceramic end cap 28 also has a generally hemispherical cavity 36 disposed therein, radially adjacent the cavity 34.
  • a generally spherical ceramic member 38 is disposed between the two outer hemispherical cavities 34 and 36, and supportively maintained therebetween.
  • the radially inner end of each ceramic blade 24 has a generally hemispherical cavity 40 disposed therein.
  • Each inner ceramic end cap 26 also has a generally hemispherical cavity 42 disposed therein, radially adjacent the cavity 40 in the radially inner end of the blade 24.
  • a generally spherical ceramic member 44 is disposed between the two inner hemispherical cavities 40 and 42, and supportively maintained therebetween.
  • the arrangement of the radially inner cavities 40 and 42 and their respective ceramic spheres 44 generally form the points of a triangle, the sides of which are indicated by dotted lines and the letters A, B and C on FIG. 2.
  • a similar non-linear arrangement of the radially outer cavities 34 and 36 is also shown in FIG. 2, labeled A,, B, and C,.
  • the radially inner and outer spherical interlock arrangements 25 each form a tetrahedron with spherical pivots 27 and 27, respectively.
  • the radially inner and outer tetrahedrons, defined by A B C D and A, B, C, D respectively, are each loaded in compression, due to the action of biasing member 23, and are able to pivot slightly about each radially extreme vertex, D and D.
  • each vane segment 18 permits the use of spherical interfaces in a generally triangular arrangement as stated above.
  • This triangular arrangement permits each vane segment 18 to bend unrestrained about any axis. Collapse of the three blades 24 in the vane segment 18 in the axial or circumferential direction cannot occur, since this would require a parallel displacement by the two planes defined by A B C and A, B, C,, which however, is prevented by the fixed plunger biasing arrangement 23, mentioned earlier, to help keep each entire vane segment 18 in compression. Collapse by relative rotation of the two tetrahedrons defined by A B C D and A, B, C, D, cannot occur since it would be stopped by compression of any of the three blades 24.
  • FIGS. 3 and 4 The relationship of the axial positioning of the spherical interfaces for the individual blades 24, is shown in FIGS. 3 and 4.
  • FIG. 3 showing the middle blade 24 in each vane segment 18, and
  • FIG. 4 showing an end blade 24 in each three blade vane segments.
  • the non-linearity of the middle blade spherical interface 25 with respect to the end blade spherical interfaces 25 is shown in FIG. 1, and can be seen by comparing the axial displacement of the spherical interfaces 25 of FIGS. 3 and 4.
  • Each pivotable assembly 27 or 27, or radially extreme inner or outer vertex is a spherical interface, one part of each pivotable assembly being on the supportive member 30 and 32, or load plate, the other part being disposed on the outer or inner housing rings 39 and 37 respectively.
  • the supportive members, 30 and 32 include hemispherical tenons 31 which pivotably mate with hemispherical cavities 33.
  • the tenons 31 and cavities 33 comprise part of the spherical pivot assemblies 27 and 27.
  • FIG. 5 An alternative embodiment of the spherical interface is shown in FIG. 5.
  • a single ceramic airfoil blade 46 from a vane segment 18 is shown disposed between an inner ceramic end cap 48, and an outer ceramic end cap 50.
  • the load plates or supportive members 30 and 32 are similar to the earlier shown embodiment.
  • Each ceramic end cap 48 and 50 has a generally hemispherical cavity 52 disposed thereinv
  • the ceramic airfoil blade 46 has a generally hemispherical tenon 54 on both the radially inner and outer ends. disposed radially adjacent each cavity 52 in the end caps 48 and 50.
  • the tenon 54 mates with each cavity 52 in their respective end caps 48 and 50.
  • the non-linear or generally triangular arrangement of the spherical interfaces between the blades 46 and end caps 48 and 50 would be the same as that shown in FIGS. 1 and 2.
  • a biasing means, 23, as shown in FIGS. 1 and 2 would provide a compressive force on each vane segment 18 having either the independent ceramic spheres 38 and 44, between the blade 24 and end caps 26 and 28, or the embodiment using a tenon 54 mating with cavities 52 in the end caps 48 and 50.
  • One of the tensons 54 shown in FIG. 5, could be a ceramic sphere.
  • the support arrangement would then be a combination of the embodiments of FIGS. 4 and 5, as shown in FIG. 6a.
  • FIG. 6 utilizes the disposition of a tenon 56 on each end cap 58 itself, mating with hemispherical cavities 60 on each end of an airfoil blade 62.
  • FIG. 6a utilizes a tenon 72 and hemispherical cavity 74 arrangement between one end of a blade 75 and end cap 76, and a ceramic sphere 78 and two hemispherical cavities 79 and 81 in the blade 75 and other associated end cap 80.
  • a ceramic stop 64 may be used.
  • the ceramic stop 64 is a sphere having a quadrant removed.
  • the stop 64 is restrained in a cavity 66 in an inner or outer end cap 68.
  • the stop 64 being bonded to the end cap 68.
  • a blade 69, before excessive twisting, will come into contact with wall portions of the stop 64, and prevent failure of the blade 69 or collapse of the vane segment 18.
  • the use of vane segments 18 described, in any case, require a different number of those vane segments 18 in a nozzle arrangement 12 than are usually present in the prior art. Additionally, the use of more support members 30 and 32, eliminate troublesome leakage paths, dangerous harmonics, and would reduce the number of seals required. It is understood that each alternative support embodiment utilizes the generally tetrahedronal pattern of support points in the assembled vane segment, and the triangular pattern of support points between the blades and their end caps.
  • the contact surfaces of the spherical interfaces have the advantage that they permit rotation in three directions, of each blade.
  • the rotation may be caused by: compressive spring loading a normal load; gas loading tangential load; and gas twisting moment a twist load.
  • An inlet nozzle for a gas turbine having a radially inner shroud ring, a radially outer shroud ring coaxial with said inner shroud ring, and an annular array of vane segments compressively retained therebetween, each of said vane segments comprising:
  • a first support means extending across said outer arcuate segment for retaining said outer end caps in proper orientation, with each end cap in opposed facing relationship to the radially outer end of a blade;
  • a second support means extending across said inner arcuate segment for retaining said inner end caps in proper orientation, with each end cap in opposed facing relationship to the radially inner end of a blade;
  • h. means providing a ball and socket engagement between each end cap and the adjacent radial end of i the blade associated therewith;
  • said vertices of said first included triangle in conjunction with said ball and socket engagement of said outer shroud to said support means define a substantially stable tetrahedral relationship between the points of engagement for the radially outer support of each vane segment;
  • said vertices of said second included triangle in conjunction with said ball and socket engagement of said inner shroud to said inner support means define substantially stable tetrahedral relationship between the points of engagement for the radially inner support of each vane segment.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US472753A 1974-05-23 1974-05-23 Gas turbine inlet vane structure utilizing a stable ceramic spherical interface arrangement Expired - Lifetime US3910716A (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
US472753A US3910716A (en) 1974-05-23 1974-05-23 Gas turbine inlet vane structure utilizing a stable ceramic spherical interface arrangement
CA225,297A CA1001953A (en) 1974-05-23 1975-04-23 Gas turbine inlet vane structure utilizing a stable ceramic spherical interface arrangement
IT23577/75A IT1038328B (it) 1974-05-23 1975-05-21 Struttura di palette d ingresso di una turbina a gas che impiega una disposizione di contatto stabile con elemento sferico in ceramica
JP6111575A JPS5335205B2 (it) 1974-05-23 1975-05-23

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Cited By (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4076451A (en) * 1976-03-05 1978-02-28 United Technologies Corporation Ceramic turbine stator
FR2456836A1 (fr) * 1979-05-18 1980-12-12 Avco Corp Tuyere composite en metal et ceramique, pour turbine
FR2467980A1 (fr) * 1979-10-22 1981-04-30 Gen Electric Coussin resistant a haute temperature pour piece de faible ductilite
US4302149A (en) * 1980-02-19 1981-11-24 General Motors Corporation Ceramic vane drive joint
US4643636A (en) * 1985-07-22 1987-02-17 Avco Corporation Ceramic nozzle assembly for gas turbine engine
US4768924A (en) * 1986-07-22 1988-09-06 Pratt & Whitney Canada Inc. Ceramic stator vane assembly
US4832568A (en) * 1982-02-26 1989-05-23 General Electric Company Turbomachine airfoil mounting assembly
US4861229A (en) * 1987-11-16 1989-08-29 Williams International Corporation Ceramic-matrix composite nozzle assembly for a turbine engine
EP0355312A1 (de) * 1988-08-03 1990-02-28 Asea Brown Boveri Ag Axialdurchströmte Turbine mit radial-axialer erster Stufe
US5074752A (en) * 1990-08-06 1991-12-24 General Electric Company Gas turbine outlet guide vane mounting assembly
US5141395A (en) * 1991-09-05 1992-08-25 General Electric Company Flow activated flowpath liner seal
US6648597B1 (en) 2002-05-31 2003-11-18 Siemens Westinghouse Power Corporation Ceramic matrix composite turbine vane
US20040043889A1 (en) * 2002-05-31 2004-03-04 Siemens Westinghouse Power Corporation Strain tolerant aggregate material
EP1582698A1 (en) * 2004-03-31 2005-10-05 General Electric Company Integral covered nozzle with attached overcover
US20050254942A1 (en) * 2002-09-17 2005-11-17 Siemens Westinghouse Power Corporation Method of joining ceramic parts and articles so formed
US7093359B2 (en) 2002-09-17 2006-08-22 Siemens Westinghouse Power Corporation Composite structure formed by CMC-on-insulation process
US20080181766A1 (en) * 2005-01-18 2008-07-31 Siemens Westinghouse Power Corporation Ceramic matrix composite vane with chordwise stiffener
US20110041313A1 (en) * 2009-08-24 2011-02-24 James Allister W Joining Mechanism with Stem Tension and Interlocked Compression Ring
US20130216361A1 (en) * 2012-02-22 2013-08-22 Propheter-Hinckley Tracy A Vane assembly for a gas turbine engine
CN103321692A (zh) * 2012-03-20 2013-09-25 通用电气公司 隔热装置
US20150016972A1 (en) * 2013-03-14 2015-01-15 Rolls-Royce North American Technologies, Inc. Bi-cast turbine vane
US9279335B2 (en) 2011-08-03 2016-03-08 United Technologies Corporation Vane assembly for a gas turbine engine
US9492780B2 (en) 2014-01-16 2016-11-15 Bha Altair, Llc Gas turbine inlet gas phase contaminant removal
EP3315729A1 (de) * 2016-10-26 2018-05-02 MTU Aero Engines GmbH Ellipsoidische innere leitschaufellagerung
US10260362B2 (en) 2017-05-30 2019-04-16 Rolls-Royce Corporation Turbine vane assembly with ceramic matrix composite airfoil and friction fit metallic attachment features
US10502136B2 (en) 2014-10-06 2019-12-10 Bha Altair, Llc Filtration system for use in a gas turbine engine assembly and method of assembling thereof
EP3988767A1 (de) * 2020-10-21 2022-04-27 3BE Berliner Beratungs- und Beteiligungs- Gesellschaft mbH Radialgasturbine mit stützlager

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5829308Y2 (ja) * 1977-01-22 1983-06-27 日信工業株式会社 二輪車用デイスクブレ−キにおける固定摩擦パツドの取付装置
JPS59177702U (ja) * 1984-04-12 1984-11-28 三菱重工業株式会社 ラジアル型タ−ビンノズル

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3066911A (en) * 1959-05-12 1962-12-04 Thompson Ramo Wooldridge Inc Nozzle and turbine wheel shroud support
US3843279A (en) * 1972-06-21 1974-10-22 Rolls Royce 1971 Ltd Stator assembly for gas turbine engines which accommodate circumferential and axial expansion of engine components

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3066911A (en) * 1959-05-12 1962-12-04 Thompson Ramo Wooldridge Inc Nozzle and turbine wheel shroud support
US3843279A (en) * 1972-06-21 1974-10-22 Rolls Royce 1971 Ltd Stator assembly for gas turbine engines which accommodate circumferential and axial expansion of engine components

Cited By (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4076451A (en) * 1976-03-05 1978-02-28 United Technologies Corporation Ceramic turbine stator
FR2456836A1 (fr) * 1979-05-18 1980-12-12 Avco Corp Tuyere composite en metal et ceramique, pour turbine
FR2467980A1 (fr) * 1979-10-22 1981-04-30 Gen Electric Coussin resistant a haute temperature pour piece de faible ductilite
US4312599A (en) * 1979-10-22 1982-01-26 General Electric Company High temperature article, article retainer, and cushion
US4302149A (en) * 1980-02-19 1981-11-24 General Motors Corporation Ceramic vane drive joint
US4832568A (en) * 1982-02-26 1989-05-23 General Electric Company Turbomachine airfoil mounting assembly
US4643636A (en) * 1985-07-22 1987-02-17 Avco Corporation Ceramic nozzle assembly for gas turbine engine
US4768924A (en) * 1986-07-22 1988-09-06 Pratt & Whitney Canada Inc. Ceramic stator vane assembly
US4861229A (en) * 1987-11-16 1989-08-29 Williams International Corporation Ceramic-matrix composite nozzle assembly for a turbine engine
EP0355312A1 (de) * 1988-08-03 1990-02-28 Asea Brown Boveri Ag Axialdurchströmte Turbine mit radial-axialer erster Stufe
US4948333A (en) * 1988-08-03 1990-08-14 Asea Brown Boveri Ltd. Axial-flow turbine with a radial/axial first stage
CH676735A5 (it) * 1988-08-03 1991-02-28 Asea Brown Boveri
US5074752A (en) * 1990-08-06 1991-12-24 General Electric Company Gas turbine outlet guide vane mounting assembly
US5141395A (en) * 1991-09-05 1992-08-25 General Electric Company Flow activated flowpath liner seal
US7067447B2 (en) 2002-05-31 2006-06-27 Siemens Power Generation, Inc. Strain tolerant aggregate material
US6709230B2 (en) 2002-05-31 2004-03-23 Siemens Westinghouse Power Corporation Ceramic matrix composite gas turbine vane
US20040043889A1 (en) * 2002-05-31 2004-03-04 Siemens Westinghouse Power Corporation Strain tolerant aggregate material
US6648597B1 (en) 2002-05-31 2003-11-18 Siemens Westinghouse Power Corporation Ceramic matrix composite turbine vane
US7093359B2 (en) 2002-09-17 2006-08-22 Siemens Westinghouse Power Corporation Composite structure formed by CMC-on-insulation process
US9068464B2 (en) 2002-09-17 2015-06-30 Siemens Energy, Inc. Method of joining ceramic parts and articles so formed
US20050254942A1 (en) * 2002-09-17 2005-11-17 Siemens Westinghouse Power Corporation Method of joining ceramic parts and articles so formed
US20050220622A1 (en) * 2004-03-31 2005-10-06 General Electric Company Integral covered nozzle with attached overcover
EP1582698A1 (en) * 2004-03-31 2005-10-05 General Electric Company Integral covered nozzle with attached overcover
US20080181766A1 (en) * 2005-01-18 2008-07-31 Siemens Westinghouse Power Corporation Ceramic matrix composite vane with chordwise stiffener
US7435058B2 (en) 2005-01-18 2008-10-14 Siemens Power Generation, Inc. Ceramic matrix composite vane with chordwise stiffener
US20110041313A1 (en) * 2009-08-24 2011-02-24 James Allister W Joining Mechanism with Stem Tension and Interlocked Compression Ring
US8256088B2 (en) 2009-08-24 2012-09-04 Siemens Energy, Inc. Joining mechanism with stem tension and interlocked compression ring
US9279335B2 (en) 2011-08-03 2016-03-08 United Technologies Corporation Vane assembly for a gas turbine engine
US20130216361A1 (en) * 2012-02-22 2013-08-22 Propheter-Hinckley Tracy A Vane assembly for a gas turbine engine
US9273565B2 (en) * 2012-02-22 2016-03-01 United Technologies Corporation Vane assembly for a gas turbine engine
CN103321692A (zh) * 2012-03-20 2013-09-25 通用电气公司 隔热装置
CN103321692B (zh) * 2012-03-20 2017-04-26 通用电气公司 隔热装置
US20150016972A1 (en) * 2013-03-14 2015-01-15 Rolls-Royce North American Technologies, Inc. Bi-cast turbine vane
US9803486B2 (en) * 2013-03-14 2017-10-31 Rolls-Royce North American Technologies Inc. Bi-cast turbine vane
US10612402B2 (en) 2013-03-14 2020-04-07 Rolls-Royce North American Technologies Inc. Method of assembly of bi-cast turbine vane
US9492780B2 (en) 2014-01-16 2016-11-15 Bha Altair, Llc Gas turbine inlet gas phase contaminant removal
US10502136B2 (en) 2014-10-06 2019-12-10 Bha Altair, Llc Filtration system for use in a gas turbine engine assembly and method of assembling thereof
EP3315729A1 (de) * 2016-10-26 2018-05-02 MTU Aero Engines GmbH Ellipsoidische innere leitschaufellagerung
US10294814B2 (en) 2016-10-26 2019-05-21 MTU Aero Engines AG Ellipsoidal inner central blade storage space
US10260362B2 (en) 2017-05-30 2019-04-16 Rolls-Royce Corporation Turbine vane assembly with ceramic matrix composite airfoil and friction fit metallic attachment features
EP3988767A1 (de) * 2020-10-21 2022-04-27 3BE Berliner Beratungs- und Beteiligungs- Gesellschaft mbH Radialgasturbine mit stützlager

Also Published As

Publication number Publication date
IT1038328B (it) 1979-11-20
CA1001953A (en) 1976-12-21
JPS5335205B2 (it) 1978-09-26
JPS511820A (it) 1976-01-09

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