US3781125A - Gas turbine nozzle vane structure - Google Patents

Gas turbine nozzle vane structure Download PDF

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Publication number
US3781125A
US3781125A US00241943A US3781125DA US3781125A US 3781125 A US3781125 A US 3781125A US 00241943 A US00241943 A US 00241943A US 3781125D A US3781125D A US 3781125DA US 3781125 A US3781125 A US 3781125A
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US
United States
Prior art keywords
end wall
outer shroud
structure recited
plenum chamber
sleeve
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
US00241943A
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English (en)
Inventor
T Rahaim
L Kish
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CBS Corp
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Westinghouse Electric Corp
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Filing date
Publication date
Application filed by Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
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Publication of US3781125A publication Critical patent/US3781125A/en
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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
    • 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

Definitions

  • One of the presently employed arrangements employs an annular nozzle vane structure formed in arcuate groups with a plenum chamber in the outer shroud structure to which pressurized cooling air is admitted. The air is then directed to each of a plurality of hollow nozzle vanes in the group to prevent overheating thereof by the hot motive gases flowing therepast in operation.
  • This arrangement is more fully shown and described in A. J. Scalzo and A. Zabrodsky US. Pat. No. 3,529,903, issued Sept. 22, 1970, and assigned to the same assignee as this invention.
  • This temperature difference creates large temperature gradients in the forward and rearward edge walls of the outer shroud structure that can cause stress cracks to occur therein with attendant leakage of coolant air from the plenum chamber directly into the motive gas stream with loss of cooling effect on the nozzle vane structure.
  • a pair of openended key-hole shpaed slots are provided in the outer shroud for each vane in the arcuate nozzle vane structure.
  • One of these slots is formed in the forward arcuate wall and the other is formed in the rearward arcuate wall of the shroud to pennit thermal growth to occur without thermal internal stresses of a damaging nature.
  • Each of the slots has a tubular thin walled split sleeve received in the circular portion of the slot and attached to the shroud wall by a peripheral weld.
  • the sleeve is threaded internally and disposed with the cut in its periphery in registry with the slot.
  • the split sleeve Since the split sleeve is thin and of C-shaped crosssection, ithas little, if any, stiffening effect'on the slotted shroudwall, thereby allowing thermal growth and the accommodating change inshape to take place during operation, and preventing excessive internal thermal stresses to occur. During such thermal deformation, the mating thread portions act as a labyrinthian seal to minimize leakage of coolant air therethrough. The main sealing action, however, occurs between the bolt head and the shroud wall, since the sealing face of the bolt is in a plane parallel to that in which any deformation will occur.
  • FIG. 1 is an axial sectional view of a portion of an axial flow gas turbine having a stationary annular nozzle vane structure formed in accordance with the invention
  • FIG. 2 is an enlarged front elevational view showing one of the arcuate nozzle vane group structures
  • FIG. 3 is a fragmentary sectional view of the outer shroud showing one of the sealing structures therein;
  • FIG. 4 is a perspective view of one of the nozzle vane group structures with the corner plate rewound to show internal details.
  • FIG. I there is shown a portion of an axial flow gas turbine 10. Only the upper radial half of the turbine is shown, since the lower half is substantially identical to the upper half.
  • the turbine 10 comprises an outer casing 11 of generally tubular or annular shape, an inner casing 12 of annular shape encompassed by the outer casing 11, and a rotor 14 rotatably supported within the inner casing 12 in any suitable manner and having at least one annular row of blades 16.
  • Cooperatively associated with the rotor blades 16 to form a motive fluid expansion stage is an annular row of stationary nozzle vanes 18 supported within the inner casing 12.
  • the stationary nozzle vanes 18 are provided with a radially inwardly extending air foil shaped vane portion 19, an inner arcuate shroud portion 20 and an outer shroud portion 22 received in an annular groove 24 formed in'the inner casing 12.
  • the inner casing 12 is suitably secured to the outer casing 11 by an annular radially extending flange 26 and jointly therewith forms an annular air space 27 that surrounds the inner casing.
  • the stationary nozzle vanes 18 are of hollow form, with the outer shroud portion 22 of box-like form and defining an arcuately shaped plenum chamber 28.
  • the vane portions 19 are provided with passages 31 and a radial series of outlet orifices 32 in their trailing portions.
  • the passages 31 provide a fluid communication between the plenum chamber 28 and the orifices 31.
  • the nozzle vanes 18 are preferably formed in arcuate groups 34 (as best shown in FIGS. 2 and 4) with a plurality of the vane portions 19 connected in parallel fluid flow communication with the common outer plenum chamber 28.
  • the arcuate vane group 34 extends through a central angle of 45,with the inner and outer shrouds 20 and 22 integrally formed with four of the vanes 19, as by casting. Since each arcuate nozzle vane group 34 subtends a central angle of 45 the annular row of nozzle vanes would include eight such groups disposed in end-to-end relation, as well known in the art (example: 8 X 45 360).
  • the outer shroud portion 22 is provided with a forward end wall 35 and a rearward end wall 36 (both of arcuate shape and subtending the 45 central angle, as described above) and a pair of opposed side walls 37 to impart the box-like shape to an open cavity 38 (FIG. 4) that is employed as the plenum chamber 28.
  • An outside cover plate 40 with its peripheral portions held in a grooved rail structure 41, 42 is nested in a suitable peripheral flanged portion 44 to complete the enclosure for the plenum chamber 28.
  • the nozzle vane groups 34 are keyed in the casing groove 24 and are retained therein by a tubular fitting 46 threadedly received in the inner casing 12 and having a cylindrical portion 47 extending through a mating opening 48 in the cover plate 40.
  • the fitting 46 has a central bore 50 extending therethrough to provide a fluid communication between the intercasing space 27 and the plenum chamber 28.
  • a tube or pipe 51 received in the outer casing 11 provides pressurized coolant fluid, such as pressurized air from any suitable source (not shown) to the space 27 during operation.
  • hot motive fluid such as pressurized combustion gas generated in a suitable fuel combustion chamber (not shown) is directed through an inlet passageway 52 past the stationary nozzle vanes 18 and the rotor blades 16, in the direction indicated by the arrows 53 with resulting expansion of the motive fluid to rotate the rotor 14.
  • the motive gases directed past the nozzle vanes 18 are hottor than the vanes can withstand safely, accordingly pressurized coolant air is continuously provided to the nozzle vanes 18, as previously described, to maintain them within safe temperative limits and is then ejected through the outlet orifices 32 into the stream of hot motive fluid.
  • the forward end wall 35 of the outer shroud 22 is provided with a plurality of spaced open-ended slots 54 extending radially inwardly and terminating in circular shaped bottom portions 55.
  • the slots 54 may be considered to be of key-hole shape.
  • the key-hole slots 54 are identical and each slot extends substantially the entire radial width of the shroud wall 35 to lend curvature flexibility to the shroud 22.
  • a thin walled tubular sleeve member 57 has an annular end portion 58 of reduced diameter received in mating engagement with the circular portion 55 of the slot 54 and is split by a slit 60 into a C-shaped (see FIG. 4). The thus formed C-shaped sleeve 57 is secured to the wall 35, such as by welding W along its entire periphery, with its slit 60 in registry with the associated slot 54.
  • the sleeve 57 is threaded internally and a bolt-like threaded, male sealing member 62 having a circular head 63 at one end and a screwdriver groove 65 in the other end is threadedly received therein.
  • the head 63 is of cylindrical shape and of the same diameter as the circular slot portion 55 and has an inner planar face 65.
  • the bolt heads 63 are shown in their final machined state in solid lines, but are shown in their original contour by the dot-dash lines, and the portion P is cut away between the dot-dash lines and the final contour after installation in the outer shroud.
  • This contour provides a flush assembly with the forward shroud wall 35 and permits a close fitting in the inner casing groove 24.
  • each nozzle vane group 34 four pairs are provided in each nozzle vane group 34, one pair for each vane 19, and that the slots are substantially equally spaced along the outer periphery of the shroud wall 35.
  • the rear wall 36 of the outer shroud 22 (FIG. 4) is substantially similar to the forward wall 35 and is provided with a plurality of (in this example, seven) slots 54, each fitted with a split sleeve 57 and an associated male threaded bolt-like member 62.
  • six of the heads 63 are shown in the original contour before final machining to a flush finish with the surface of the shroud wall 36 and one of the heads 63 (at the right end of the nozzle group 34) is illustrated in the final contoured shape.
  • the bolt heads 63 are machined to a flush fit to permit a good fit with the groove 24 in the inner casing 12 (FIG. 1).
  • the side walls 70, 71 of the casing groove 24 are extended radially inwardly to overlap this portion of the slots (see FIG. 1).
  • the forward and rearward walls 35 and 36 are slotted and thus rendered flexible, they permit the shroud to undergo an accommodating change of shape without creating high local internal stresses Since the outer shroud tends to increase its radius of curvature, due to thermal growth the slots 54 tend to deform in shape and this deformation is freely permitted by the split sleeves 57, while still acting to prevent leakage therepast from the plenum chamber 28, as previously described.
  • a nozzle vane structure for a hot elastic fluid utilizing machine comprising a plurality of vane portions of generally airfoil cross-section, an inner arcuate shroud segment integrally connected to one end of said vane portions, an outer arcuate shroud segment integrally connected to the opposite end of said vane portions, and outer shroud segment having a forward end wall and a rearward end wall of arcuate cross-section, means cooperating with said outer shroud segment to form a plenum chamber, means for admitting a pressurized coolant fluid to said plenum chamber, each of said vanes having a passage therein communicating with said plenum chamber for directing coolant fluid from said chamber through said vane portions to cool the latter, at least one of said end walls having an open-ended circular slot to pennit thermal growth during operation, an internally threaded sleeve member received in said slot and attached to said one end wall, and a threaded male member disposed in mating threaded coolant therepast during distortion of said s

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US00241943A 1972-04-07 1972-04-07 Gas turbine nozzle vane structure Expired - Lifetime US3781125A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US24194372A 1972-04-07 1972-04-07

Publications (1)

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US3781125A true US3781125A (en) 1973-12-25

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US00241943A Expired - Lifetime US3781125A (en) 1972-04-07 1972-04-07 Gas turbine nozzle vane structure

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US (1) US3781125A (it)
JP (1) JPS5017603B2 (it)
CA (1) CA965352A (it)
CH (1) CH576065A5 (it)
DE (1) DE2315745A1 (it)
FR (1) FR2179455A5 (it)
GB (1) GB1373898A (it)
IT (1) IT982711B (it)
NL (1) NL7304355A (it)
SE (1) SE379079B (it)

Cited By (37)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3841787A (en) * 1973-09-05 1974-10-15 Westinghouse Electric Corp Axial flow turbine structure
US3892497A (en) * 1974-05-14 1975-07-01 Westinghouse Electric Corp Axial flow turbine stationary blade and blade ring locking arrangement
USB561712I5 (it) * 1975-03-25 1976-02-17
US3957391A (en) * 1975-03-25 1976-05-18 United Technologies Corporation Turbine cooling
US4126405A (en) * 1976-12-16 1978-11-21 General Electric Company Turbine nozzle
US4180371A (en) * 1978-03-22 1979-12-25 Avco Corporation Composite metal-ceramic turbine nozzle
FR2575224A1 (fr) * 1984-12-21 1986-06-27 United Technologies Corp Segment d'etancheite refroidissable pour une machine rotative, notamment pour un moteur a turbine a gaz a flux axial
US5071313A (en) * 1990-01-16 1991-12-10 General Electric Company Rotor blade shroud segment
US6537022B1 (en) * 2001-10-05 2003-03-25 General Electric Company Nozzle lock for gas turbine engines
US20030206799A1 (en) * 2002-05-02 2003-11-06 Scott John M. Casing section
US20030231957A1 (en) * 2002-02-22 2003-12-18 Power Technology Incorporated Compressor stator vane
US6773229B1 (en) * 2003-03-14 2004-08-10 General Electric Company Turbine nozzle having angel wing seal lands and associated welding method
US20060013685A1 (en) * 2004-07-14 2006-01-19 Ellis Charles A Vane platform rail configuration for reduced airfoil stress
US20060099078A1 (en) * 2004-02-03 2006-05-11 Honeywell International Inc., Hoop stress relief mechanism for gas turbine engines
US20070128020A1 (en) * 2005-12-05 2007-06-07 Snecma Bladed stator for a turbo-engine
US20070166154A1 (en) * 2004-07-14 2007-07-19 Power Systems Mfg., Llc Vane platform rail configuration for reduced airfoil stress
US20080148737A1 (en) * 2006-12-21 2008-06-26 Power Systems Manufacturing, Llc Turbine Static Structure for Reduced Leakage Air
US20080282541A1 (en) * 2002-02-22 2008-11-20 Anderson Rodger O Compressor stator vane
US20090110552A1 (en) * 2007-10-31 2009-04-30 Anderson Rodger O Compressor stator vane repair with pin
FR2928962A1 (fr) * 2008-03-19 2009-09-25 Snecma Sa Distributeur de turbine a pales creuses.
US20100247303A1 (en) * 2009-03-26 2010-09-30 General Electric Company Duct member based nozzle for turbine
US20110048023A1 (en) * 2009-09-02 2011-03-03 Pratt & Whitney Canada Corp. Fuel nozzle swirler assembly
US20110064580A1 (en) * 2009-09-16 2011-03-17 United Technologies Corporation Turbofan flow path trenches
US20110085894A1 (en) * 2008-05-26 2011-04-14 Alstom Technology Ltd Gas turbine with a stator blade
US8888442B2 (en) 2012-01-30 2014-11-18 Pratt & Whitney Canada Corp. Stress relieving slots for turbine vane ring
US20140341731A1 (en) * 2011-05-30 2014-11-20 Siemens Aktiengesellschaft Piston seal ring
EP2937518A1 (en) * 2014-04-21 2015-10-28 Honeywell International Inc. Gas turbine engine components having sealed stress relief slots and methods for the fabrication thereof
US20170145833A1 (en) * 2015-11-23 2017-05-25 United Technologies Corporation Baffle for a component of a gas turbine engine
US20180179898A1 (en) * 2015-09-17 2018-06-28 Safran Aircraft Engines Nozzle sector for a turbine engine with differentially cooled blades
US20180223691A1 (en) * 2017-02-03 2018-08-09 United Technologies Corporation Case flange with stress reducing features
EP3428402A1 (de) * 2017-07-11 2019-01-16 MTU Aero Engines GmbH Leitschaufelsegment mit gekrümmter entlastungsfuge
US20200088049A1 (en) * 2018-09-18 2020-03-19 General Electric Company Airfoil shroud assembly using tenon with externally threaded stud and nut
US10822980B2 (en) 2013-04-11 2020-11-03 Raytheon Technologies Corporation Gas turbine engine stress isolation scallop
US11111804B2 (en) * 2019-03-11 2021-09-07 Raytheon Technologies Corporation Inserts for slotted integrally bladed rotor
US11459900B2 (en) * 2020-06-16 2022-10-04 Toshiba Energy Systems & Solutions Corporation Turbine stator blade
US11814991B1 (en) 2022-07-28 2023-11-14 General Electric Company Turbine nozzle assembly with stress relief structure for mounting rail
US11885241B1 (en) * 2022-07-28 2024-01-30 General Electric Company Turbine nozzle assembly with stress relief structure for mounting rail

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
SE411931B (sv) * 1975-03-25 1980-02-11 United Technologies Corp Anordning vid turbinmunstycken for gasturbinmotor
FR2914017B1 (fr) * 2007-03-20 2011-07-08 Snecma Dispositif d'etancheite pour un circuit de refroidissement, carter inter-turbine en etant equipe et turboreacteur les comportant
CN111594276A (zh) * 2020-05-19 2020-08-28 哈尔滨汽轮机厂有限责任公司 一种汽轮机整体式喷嘴组用外围带及其加工、装配方法

Cited By (62)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3841787A (en) * 1973-09-05 1974-10-15 Westinghouse Electric Corp Axial flow turbine structure
US3892497A (en) * 1974-05-14 1975-07-01 Westinghouse Electric Corp Axial flow turbine stationary blade and blade ring locking arrangement
USB561712I5 (it) * 1975-03-25 1976-02-17
US3957391A (en) * 1975-03-25 1976-05-18 United Technologies Corporation Turbine cooling
US3992126A (en) * 1975-03-25 1976-11-16 United Technologies Corporation Turbine cooling
US4126405A (en) * 1976-12-16 1978-11-21 General Electric Company Turbine nozzle
US4180371A (en) * 1978-03-22 1979-12-25 Avco Corporation Composite metal-ceramic turbine nozzle
FR2575224A1 (fr) * 1984-12-21 1986-06-27 United Technologies Corp Segment d'etancheite refroidissable pour une machine rotative, notamment pour un moteur a turbine a gaz a flux axial
US4650395A (en) * 1984-12-21 1987-03-17 United Technologies Corporation Coolable seal segment for a rotary machine
US5071313A (en) * 1990-01-16 1991-12-10 General Electric Company Rotor blade shroud segment
US6537022B1 (en) * 2001-10-05 2003-03-25 General Electric Company Nozzle lock for gas turbine engines
US7984548B2 (en) 2002-02-22 2011-07-26 Drs Power Technology Inc. Method for modifying a compressor stator vane
US20030231957A1 (en) * 2002-02-22 2003-12-18 Power Technology Incorporated Compressor stator vane
US20080282541A1 (en) * 2002-02-22 2008-11-20 Anderson Rodger O Compressor stator vane
US6984108B2 (en) * 2002-02-22 2006-01-10 Drs Power Technology Inc. Compressor stator vane
US6991427B2 (en) * 2002-05-02 2006-01-31 Rolls-Royce Plc Casing section
US20030206799A1 (en) * 2002-05-02 2003-11-06 Scott John M. Casing section
US6773229B1 (en) * 2003-03-14 2004-08-10 General Electric Company Turbine nozzle having angel wing seal lands and associated welding method
US20060099078A1 (en) * 2004-02-03 2006-05-11 Honeywell International Inc., Hoop stress relief mechanism for gas turbine engines
US7097422B2 (en) 2004-02-03 2006-08-29 Honeywell International, Inc. Hoop stress relief mechanism for gas turbine engines
US20070166154A1 (en) * 2004-07-14 2007-07-19 Power Systems Mfg., Llc Vane platform rail configuration for reduced airfoil stress
US7229245B2 (en) * 2004-07-14 2007-06-12 Power Systems Mfg., Llc Vane platform rail configuration for reduced airfoil stress
US7293957B2 (en) * 2004-07-14 2007-11-13 Power Systems Mfg., Llc Vane platform rail configuration for reduced airfoil stress
US20060013685A1 (en) * 2004-07-14 2006-01-19 Ellis Charles A Vane platform rail configuration for reduced airfoil stress
US7780398B2 (en) 2005-12-05 2010-08-24 Snecma Bladed stator for a turbo-engine
FR2894282A1 (fr) * 2005-12-05 2007-06-08 Snecma Sa Distributeur de turbine de turbomachine ameliore
EP1793093A3 (fr) * 2005-12-05 2008-12-03 Snecma Distributeur de turbine de turbomachine amélioré
CN1978870B (zh) * 2005-12-05 2012-05-30 斯奈克玛 涡轮发动机改良型带叶片定子
US20070128020A1 (en) * 2005-12-05 2007-06-07 Snecma Bladed stator for a turbo-engine
US20080148737A1 (en) * 2006-12-21 2008-06-26 Power Systems Manufacturing, Llc Turbine Static Structure for Reduced Leakage Air
US7958735B2 (en) 2006-12-21 2011-06-14 Power Systems Manufacturing, Llc Turbine static structure for reduced leakage air
US20090110552A1 (en) * 2007-10-31 2009-04-30 Anderson Rodger O Compressor stator vane repair with pin
FR2928962A1 (fr) * 2008-03-19 2009-09-25 Snecma Sa Distributeur de turbine a pales creuses.
US8210797B2 (en) * 2008-05-26 2012-07-03 Alstom Technology Ltd Gas turbine with a stator blade
US20110085894A1 (en) * 2008-05-26 2011-04-14 Alstom Technology Ltd Gas turbine with a stator blade
US20100247303A1 (en) * 2009-03-26 2010-09-30 General Electric Company Duct member based nozzle for turbine
US8371810B2 (en) 2009-03-26 2013-02-12 General Electric Company Duct member based nozzle for turbine
US20110048023A1 (en) * 2009-09-02 2011-03-03 Pratt & Whitney Canada Corp. Fuel nozzle swirler assembly
US8555649B2 (en) 2009-09-02 2013-10-15 Pratt & Whitney Canada Corp. Fuel nozzle swirler assembly
US8403645B2 (en) 2009-09-16 2013-03-26 United Technologies Corporation Turbofan flow path trenches
US8834129B2 (en) 2009-09-16 2014-09-16 United Technologies Corporation Turbofan flow path trenches
US20110064580A1 (en) * 2009-09-16 2011-03-17 United Technologies Corporation Turbofan flow path trenches
US9422823B2 (en) * 2011-05-30 2016-08-23 Siemens Aktiengesellschaft Piston seal ring
US20140341731A1 (en) * 2011-05-30 2014-11-20 Siemens Aktiengesellschaft Piston seal ring
US8888442B2 (en) 2012-01-30 2014-11-18 Pratt & Whitney Canada Corp. Stress relieving slots for turbine vane ring
US10822980B2 (en) 2013-04-11 2020-11-03 Raytheon Technologies Corporation Gas turbine engine stress isolation scallop
EP2937518A1 (en) * 2014-04-21 2015-10-28 Honeywell International Inc. Gas turbine engine components having sealed stress relief slots and methods for the fabrication thereof
US9506365B2 (en) 2014-04-21 2016-11-29 Honeywell International Inc. Gas turbine engine components having sealed stress relief slots and methods for the fabrication thereof
US20180179898A1 (en) * 2015-09-17 2018-06-28 Safran Aircraft Engines Nozzle sector for a turbine engine with differentially cooled blades
US20170145833A1 (en) * 2015-11-23 2017-05-25 United Technologies Corporation Baffle for a component of a gas turbine engine
US10370979B2 (en) * 2015-11-23 2019-08-06 United Technologies Corporation Baffle for a component of a gas turbine engine
US11035236B2 (en) 2015-11-23 2021-06-15 Raytheon Technologies Corporation Baffle for a component of a gas turbine engine
US20180223691A1 (en) * 2017-02-03 2018-08-09 United Technologies Corporation Case flange with stress reducing features
EP3428402A1 (de) * 2017-07-11 2019-01-16 MTU Aero Engines GmbH Leitschaufelsegment mit gekrümmter entlastungsfuge
US10731489B2 (en) 2017-07-11 2020-08-04 MTU Aero Engines AG Guide vane segment with curved relief gap
US11028709B2 (en) * 2018-09-18 2021-06-08 General Electric Company Airfoil shroud assembly using tenon with externally threaded stud and nut
US20200088049A1 (en) * 2018-09-18 2020-03-19 General Electric Company Airfoil shroud assembly using tenon with externally threaded stud and nut
US11111804B2 (en) * 2019-03-11 2021-09-07 Raytheon Technologies Corporation Inserts for slotted integrally bladed rotor
US11459900B2 (en) * 2020-06-16 2022-10-04 Toshiba Energy Systems & Solutions Corporation Turbine stator blade
US11814991B1 (en) 2022-07-28 2023-11-14 General Electric Company Turbine nozzle assembly with stress relief structure for mounting rail
US11885241B1 (en) * 2022-07-28 2024-01-30 General Electric Company Turbine nozzle assembly with stress relief structure for mounting rail
US20240035388A1 (en) * 2022-07-28 2024-02-01 General Electric Company Turbine nozzle assembly with stress relief structure for mounting rail

Also Published As

Publication number Publication date
FR2179455A5 (it) 1973-11-16
IT982711B (it) 1974-10-21
NL7304355A (it) 1973-10-09
JPS498609A (it) 1974-01-25
CA965352A (en) 1975-04-01
DE2315745A1 (de) 1973-10-18
GB1373898A (en) 1974-11-13
SE379079B (it) 1975-09-22
CH576065A5 (it) 1976-05-31
JPS5017603B2 (it) 1975-06-23

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