US3708242A - Supporting structure for the blades of turbomachines - Google Patents

Supporting structure for the blades of turbomachines Download PDF

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Publication number
US3708242A
US3708242A US00093655A US3708242DA US3708242A US 3708242 A US3708242 A US 3708242A US 00093655 A US00093655 A US 00093655A US 3708242D A US3708242D A US 3708242DA US 3708242 A US3708242 A US 3708242A
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United States
Prior art keywords
corrugated
corrugations
elements
turbomachine according
general direction
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Expired - Lifetime
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US00093655A
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English (en)
Inventor
H Bruneau
G Langner
M Tournere
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Safran Aircraft Engines SAS
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SNECMA SAS
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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
    • F01D9/044—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators permanently, e.g. by welding, brazing, casting or the like
    • 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
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention relates to a supporting structure for blades, for example, stator blades, in a fluid handling machine having a relatively rotatable stator and rotor, hereinafter referred to as a turbomachine, such as a compressor, a turbine or a pump.
  • a turbomachine such as a compressor, a turbine or a pump.
  • a structure of this kind is often designed to withstand the key stresses, in particular bending, to which the machine is subjected and must, consequently, have a high stiffness.
  • Known structures employed for this purpose generally comprise an external structure made up, for example, of a set of cylindrical or frusto-conical casings assembled together axially, or again two half-shells joined to one another along an axial plane.
  • the stator blades are fixed, at one of their ends, to said external structure, in order to form one or more successive blade rows. It is often advantageous, where the attainment of rigidity in the assembly is concerned, to make this a rigid attachment, for example by effecting the same by brazing or welding, so that it consequently represents an embedded or restrained attachment. Equally, it is possible to provide an internal structure to which the blades are fixed by their other ends.
  • the supporting structures thus far known exhibit the drawback, in particular in the case where the blades are fixed to them in embedded fashion, that they do not provide adequate damping of the vibrations to which the blades are subjected in operation and this makes for increased risk of fatigue failure on the part of the blades.
  • These structures are, furthermore, relatively heavy and this is a particular drawback where the turbomachine is designed for part of a gas-turbine power plant for aviation applications.
  • the object of the invention is simultaneously to overcome these two drawbacks.
  • a bladesupporting structure is designed to comprise at least one corrugated or undulating annular element disposed coaxially relative to the axis of the machine, said element being formed with openings in which each of the blades is fitted at one of its ends, each of said openings extending in a direction oblique to the general directions of the folds or corrugations in said corrugated element and passing through at least two successive corrugations.
  • said corrugated element will be located between two skins or liners, so as to form the core of a sandwich structure.
  • This structure in the application with which the invention is concerned, has the essential advantage of combining two apparently contradictory qualities, namely a high structural rigidity, thus enabling the supporting structure which it forms to withstand at least some of the main stresses to which the machine is subjected, and, on the other hand, a local flexibility which makes it possible to damp the vibrations developing in the blades which are fixed to said structure.
  • This structure is, furthermore, much lighter than the conventional supporting structures whilst having the same or even better stiffness.
  • the corrugated element doing duty as the core of the structure can be assembled so that the general direction of its folds or corrugations is substantially parallel to the axis of the turbomachine, this, in particular, making for high transverse stiffness. Equally, how ever, it can be assembled in such a manner, and in particular in the case where the supporting structure does not have to withstand any substantial bending stresses, that the general direction of its folds or corrugations is substantially circumferential.
  • the core of the sandwich structure may comprise at least two generally concentric corrugated elements, whose folds or corrugations extend either in the same direction (advantageously with different or with a certain difference in phase), or in two mutually perpendicular directions. In this manner, simultaneously both the stiffness of the supporting structure and its capacity to damp blade vibrations, are improved.
  • the supporting sandwich structure can be fixed to the interior of a conventional rigid external housing casing.
  • FIG. 1 is a schematic view, in section on an axial plane, of a part of the stator of a turbomachine comprising a stator blade supporting structure in accordance with the invention
  • FIGS. 2 to 11 are schematic views relating to different embodiments of said supporting structure or to its assembly;
  • FIGS. la, 2a and 5a are sectional views on the lines Ia-la, Ila-Ila and Va-Va respectively, of the structures respectively illustrated in FIGS. 1,2 and 5;
  • FIGS. l2, l3 and 14 are explanatory diagrams relating to the vibratory phenomena occurring in the turbomachine during operation.
  • FIGS. 1 and la part of the stator of a turbomachine, such as an axial-flow compressor for a turbojet engine, has been shown, and indeed, more specifically, a supporting structure for a row of stator blades 1.
  • This structure comprises an external casing section 2 and a internal casing section 3 each constituted, in part at least, by a body of sandwich structure.
  • Each of these sections comprises an annular corrugated element 2a or 30 forming the core, between two skins" 3b, 3c or 3!), Be.
  • the general direction of the folds or corrugations in the elements 2a and 3a is substantially circumferential.
  • a ring 4 formed with a groove 4a or with a tongue so that sealed attachment to an adjacent complementary ring, belonging to the next casing section, can be effected.
  • the sandwich structure of the internal casings 3 is likewise pinched at its edges and bent in order to enable the attachment, for example by welding or brazing, of sealing rings 5. These latter operate in a manner known per se with labyrinth arrangements carried by the rotor of the machine and shown in FIG. 11.
  • the external casing section 2 and the internal casing section 3 are punched to fit the shape and setting angle of the blades 1.
  • elongated openings are formed each of which extends lengthwise in a direction oblique to the general direction of the corrugations in the corrugated elements 20 and 3a and passes through or across at least two successive corrugations.
  • the respective ends of the blades are fitted either by simply press-fitting them in position, or by securing them by some other method such as welding, brazing, diffusion or bonding.
  • H05. 2 and 2a relate to a variant embodiment in accordance with which the general direction of the folds or corrugations in the corrugated elements 20, 3a, is substantially parallel to the axis of the turbomachine.
  • the rings 4 of FIG. 1 are replaced by flanges 6 attached, for example by welding, to the two previously pinched ends of the casing section 2.
  • the adjacent flanges of two successive casing sections are advantageously assembled together by bolts.
  • FIG. 3 relates to a variant embodiment in accordance with which the external supporting structure comprises a plurality of easing sections 2, cylindrical or frusto-conical in form, assembled together axially. At least at one of its ends, the structure terminates in an attached flange 7 similar (with the exception of a shoulder 70 to the flanges 6 described in respect of FIG. 2.
  • the fixing together of the casing sections 2, on the other hand, is effected by flanges 8 which are produced by pinching and folding the sandwich structure itself, said flanges being bolted together at 9.
  • FIG. 4 illustrates a variant embodiment which differs from the arrangement shown in FIG. 1 simply by the fact that the tongued and grooved rings 4 are replaced by rings 10 with interlocking teeth 11 so that the casing sections are centered and prevented from rotating in relation to one another.
  • FIGS. 5 and 5a relate to a variant embodiment of the arrangement shown in FIGS. 2 and 2a, in accordance with which the folds or corrugations in the corrugated elements 2a and 3a extend respectively in circumferential and axial directions.
  • FIG. 5 likewise illustrates the fixing together of the successive flanges, by bolts 12.
  • FIGS. 6 and 7 relate to variant embodiments of the invention in which the core of the sandwich structure is made up of at least two concentric corrugated annular elements 120, 22a, separated from one another by an intermediate skin" 2d.
  • the folds or corrugations in said elements extend in two mutually perpendicular directions, namely circumferentially in the case of the element 12a and axially in the case of the element 22a.
  • these folds extend in the same direction, for example circumferentially.
  • the spacing or the phase of the corrugations 120 will advantageously be made different to that of the corrugations 220, this increasing the transverse rigidity of the supporting structure assembly.
  • FIG. 8 illustrates a variant embodiment of the invention in accordance with which the core of the sandwich structure comprises three concentric corrugated annular elements 102a 112a, 122a, separated from one another by intermediate skins 2e, 2f.
  • the spacing of the corrugations 112a differs from that of those 102a and 122a, and these latter can themselves be so arranged in relation to one another as to give a certain difference in phase which is beneficial in increasing the stiffness of the supporting structure assembly.
  • FIG. 9 relates to a variant embodiment of the invention in accordance with which the corrugations of the element 2a, instead of being simple as in the preceding cases, have a complex periodic profile.
  • FIGS. 10 and I1 relate to two variant embodiments of the invention in accordance with which the supporting structures 2 (sandwich structures), for the stator blades, are fixed to the interior of a conventional rigid housing constituted, for example, by a stack of rings such as those marked 20.
  • the supporting structures 2 present flanges 21, 22 fixed by bolts 23 to the housing rings 20.
  • the supporting structures 2 are fixed to the housing on the one hand in the circumferential direction by a system of splines 24, and on the other hand, in the axial direction, by the axial stacking of said structures and the fitting of retainer rings 25.
  • This figure also shows the relative arrangement of two rows of stator blades 1 and a row of rotor blades 26 carried by a disc 27. Assembled on the latter there are labyrinth arrangements 28 cooperating with the sealing rings 5.
  • FIGS. l2, l3 and 14 provide a highly schematic illustration of the vibratory phenomena to which reference has been made hereinbefore.
  • stator blades 1 exposed to vibrations of aerodynamic nature in the fluid flowing through the turbomachine, themselves experience mechanical vibrations for example in the fundamental mode 1a or at a harmonic, for example the second harmonic lb.
  • the amplitudes of these vibrations, illustrated in FIG. 12, have been exaggerated simply in order to clarify the drawing.
  • FIGS. 13 and 14 show, the elongations e are along the large axis B-B of the blade profile.
  • FIG. I3 corresponds to the case in which the corrugations in the material are axially disposed; this material then has a high radial rigidity and, by contrast, a certain degree of circumferential flexibility.
  • FIG. 14 relates to the case in which the corrugations in the material extend circumferentially; its rigidity and flexibility are then respectively circumferential and axial.
  • the main stresses to which the turbomachine is subjected being generally axial and radial, it may be a useful feature to use for the external supporting structure a sandwich material incorporating at least one element with axial corrugations, as shown in FIGS. 2 and 6, at any rate in the case where it is necessary to take into account the internal arrangement of the corrugations in order to achieve a stiff stator assembly.
  • a sandwich material incorporating at least one element with axial corrugations as shown in FIGS. 2 and 6, at any rate in the case where it is necessary to take into account the internal arrangement of the corrugations in order to achieve a stiff stator assembly.
  • stiffness is provided anyway by a conventional external casing, as shown in FIGS. and 11, the choice of the type of corrugations depends essentially upon its anti-vibration properties.
  • stator blades supporting structures which have been illustrated in the form of casing sections designed for axial assembly, could be replaced by half-shells or shell segments assembled together along axial lines.
  • a turbomachine comprising at least one corrugated annular support element which extends coaxially with the axis of the machine, said corrugated element being perforated with a plurality of elongated openings each of which extends lengthwise in a direction oblique to the general direction of the corrugations in said corrugated element and passes through at least two successive corrugations, and a plurality of blades each having one of its ends shaped to and fixedly fitted in a corresponding elongated opening and secured to a plurality of said corrugations.
  • a turbomachine according to claim 1 further com prising two skins between which said corrugated element is located, whereby to form a sandwich structure.
  • a turbomachine according to claim I comprising two corrugated annular elements arranged in radially spaced concentric coaxial relation to the axis of the machine, each such element being perforated with a plurality of elongated openings, each opening in a corrugated element extending lengthwise in a direction oblique to the general direction of the corrugations in said corrugated element and passing through at least two corrugations in said element, each opening in one of said corrugated elements being in radial registration with an opening in the other corrugated element; and a plurality of blades each having its opposite ends fitting in a corresponding pair of radially spaced registering openings.
  • corrugations in the two respective corrugated elements are each generally of the shape of a regular sinuous curve and the curves of the two elements are out of phase.
  • annular corrugated element comprises at least two annular sections arranged in axial succeeding relation.
  • a turbomachine according to claim 1 further comprising a rigid external housing, and means fixing said annular corrugated element to the interior of said external housing.
  • a turbomachine according to claim 1 comprising at least two corrugated annular elements arranged in radially adjacent concentric coaxial relation to the axis of the machine, each such element being perforated with a plurality of elongated openings, each opening in a corrugated element extending lengthwise in a direction oblique to the general direction of the corrugations in said corrugated element and passing through at least two corrugations in said element, each opening in one of said corrugated elements being in radial registration with an opening in at least one other corrugated element; and a plurality of blades each having one of its ends fitting into at least two corresponding registering openings in such radially adjacent corrugated elements.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US00093655A 1969-12-01 1970-11-30 Supporting structure for the blades of turbomachines Expired - Lifetime US3708242A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR6941393A FR2073239A1 (enrdf_load_stackoverflow) 1969-12-01 1969-12-01

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US3708242A true US3708242A (en) 1973-01-02

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US (1) US3708242A (enrdf_load_stackoverflow)
FR (1) FR2073239A1 (enrdf_load_stackoverflow)
GB (1) GB1298868A (enrdf_load_stackoverflow)

Cited By (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3778185A (en) * 1972-08-28 1973-12-11 United Aircraft Corp Composite strut joint construction
US3836282A (en) * 1973-03-28 1974-09-17 United Aircraft Corp Stator vane support and construction thereof
US3887299A (en) * 1973-08-28 1975-06-03 Us Air Force Non-abradable turbine seal
US3924964A (en) * 1974-12-23 1975-12-09 Trane Co Axial flow fan apparatus
US4288202A (en) * 1977-06-20 1981-09-08 Aerojet-General Corporation Fan having internal polyhedral strut frame
US4305696A (en) * 1979-03-14 1981-12-15 Rolls-Royce Limited Stator vane assembly for a gas turbine engine
US4452564A (en) * 1981-11-09 1984-06-05 The Garrett Corporation Stator vane assembly and associated methods
US4695225A (en) * 1983-08-30 1987-09-22 Bbc Brown, Boveri & Company, Limited Axial swirl body for generating rotary flows
US4940386A (en) * 1987-02-05 1990-07-10 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Multiple flow turbojet engine with an outer ring of the fan outlet shrunk onto the case
US5083900A (en) * 1989-11-15 1992-01-28 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbomachine stator element
US5249418A (en) * 1991-09-16 1993-10-05 General Electric Company Gas turbine engine polygonal structural frame with axially curved panels
US5411368A (en) * 1993-11-08 1995-05-02 Allied-Signal Inc. Ceramic-to-metal stator vane assembly with braze
US5483792A (en) * 1993-05-05 1996-01-16 General Electric Company Turbine frame stiffening rails
US5494404A (en) * 1993-12-22 1996-02-27 Alliedsignal Inc. Insertable stator vane assembly
US5704762A (en) * 1993-11-08 1998-01-06 Alliedsignal Inc. Ceramic-to-metal stator vane assembly
US5823739A (en) * 1996-07-03 1998-10-20 United Technologies Corporation Containment case for a turbine engine
US6000906A (en) * 1997-09-12 1999-12-14 Alliedsignal Inc. Ceramic airfoil
US20050129514A1 (en) * 2003-06-30 2005-06-16 Snecma Moteurs Nozzle ring with adhesive bonded blading for aircraft engine compressor
US20080022692A1 (en) * 2006-07-27 2008-01-31 United Technologies Corporation Embedded mount for mid-turbine frame
DE102009007999A1 (de) * 2009-02-07 2010-08-12 Hobis Ag Leitringelement für Turbinen und Verfahren zu dessen Herstellung
WO2010094277A3 (de) * 2009-02-23 2011-06-23 Mtu Aero Engines Gmbh Gasturbinenmaschine mit einem gedämpften schaufelcluster
US20110214433A1 (en) * 2010-03-08 2011-09-08 United Technologies Corporation Strain tolerant bound structure for a gas turbine engine
US20110243742A1 (en) * 2010-03-30 2011-10-06 Snecma Stator stage for turbomachine compressor
US20140165533A1 (en) * 2012-12-18 2014-06-19 Pratt & Whitney Canada Corp. Gas turbine engine mounting ring
US20150086352A1 (en) * 2013-09-25 2015-03-26 Pratt & Whitney Canada Corp. Gas Turbine Engine Inlet Assembly and Method of Making Same
US20160025108A1 (en) * 2014-07-25 2016-01-28 Techspace Aero S.A. Vane with Sealed Lattice in a Shroud of an Axial Turbomachine Compressor
US20160024971A1 (en) * 2014-07-22 2016-01-28 Rolls-Royce Plc Vane assembly
FR3097909A1 (fr) * 2019-06-27 2021-01-01 Safran Aircraft Engines Virole interne d'un carter intermédiaire, carter intermédiaire associé avec lamelles formant amortisseurs
US11352895B2 (en) 2019-10-29 2022-06-07 Raytheon Technologies Corporation System for an improved stator assembly

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2664944B1 (fr) * 1990-07-18 1992-09-25 Snecma Compresseur forme notamment de redresseurs en couronne et procede de montage de ce compresseur.
CN112059553B (zh) * 2020-09-08 2021-11-05 中国航发沈阳黎明航空发动机有限责任公司 一种用于中介机匣的多重交叉加工方法

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA154605A (en) * 1913-11-17 1914-03-24 Alvin F. Bevard Trolley
US2724546A (en) * 1951-08-03 1955-11-22 Westinghouse Electric Corp Gas turbine apparatus
GB994568A (en) * 1964-05-07 1965-06-10 Rolls Royce Bladed structure
GB1028444A (en) * 1965-01-20 1966-05-04 Rolls Royce Compressor for a gas turbine engine
US3291382A (en) * 1964-05-08 1966-12-13 Rolls Royce Bladed structure, for example, for a gas turbine engine compressor
US3303998A (en) * 1966-07-18 1967-02-14 Gen Electric Stator casing

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA154605A (en) * 1913-11-17 1914-03-24 Alvin F. Bevard Trolley
US2724546A (en) * 1951-08-03 1955-11-22 Westinghouse Electric Corp Gas turbine apparatus
GB994568A (en) * 1964-05-07 1965-06-10 Rolls Royce Bladed structure
US3291382A (en) * 1964-05-08 1966-12-13 Rolls Royce Bladed structure, for example, for a gas turbine engine compressor
GB1028444A (en) * 1965-01-20 1966-05-04 Rolls Royce Compressor for a gas turbine engine
US3303998A (en) * 1966-07-18 1967-02-14 Gen Electric Stator casing

Cited By (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3778185A (en) * 1972-08-28 1973-12-11 United Aircraft Corp Composite strut joint construction
US3836282A (en) * 1973-03-28 1974-09-17 United Aircraft Corp Stator vane support and construction thereof
US3887299A (en) * 1973-08-28 1975-06-03 Us Air Force Non-abradable turbine seal
US3924964A (en) * 1974-12-23 1975-12-09 Trane Co Axial flow fan apparatus
US4288202A (en) * 1977-06-20 1981-09-08 Aerojet-General Corporation Fan having internal polyhedral strut frame
US4305696A (en) * 1979-03-14 1981-12-15 Rolls-Royce Limited Stator vane assembly for a gas turbine engine
US4452564A (en) * 1981-11-09 1984-06-05 The Garrett Corporation Stator vane assembly and associated methods
US4695225A (en) * 1983-08-30 1987-09-22 Bbc Brown, Boveri & Company, Limited Axial swirl body for generating rotary flows
US4940386A (en) * 1987-02-05 1990-07-10 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Multiple flow turbojet engine with an outer ring of the fan outlet shrunk onto the case
US5083900A (en) * 1989-11-15 1992-01-28 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbomachine stator element
US5249418A (en) * 1991-09-16 1993-10-05 General Electric Company Gas turbine engine polygonal structural frame with axially curved panels
GB2259551B (en) * 1991-09-16 1994-10-19 Gen Electric Gas turbine engine polygonal structural frame with axially curved panels
US5483792A (en) * 1993-05-05 1996-01-16 General Electric Company Turbine frame stiffening rails
US5411368A (en) * 1993-11-08 1995-05-02 Allied-Signal Inc. Ceramic-to-metal stator vane assembly with braze
US5704762A (en) * 1993-11-08 1998-01-06 Alliedsignal Inc. Ceramic-to-metal stator vane assembly
US5494404A (en) * 1993-12-22 1996-02-27 Alliedsignal Inc. Insertable stator vane assembly
US5823739A (en) * 1996-07-03 1998-10-20 United Technologies Corporation Containment case for a turbine engine
US6000906A (en) * 1997-09-12 1999-12-14 Alliedsignal Inc. Ceramic airfoil
US20050129514A1 (en) * 2003-06-30 2005-06-16 Snecma Moteurs Nozzle ring with adhesive bonded blading for aircraft engine compressor
US7147434B2 (en) * 2003-06-30 2006-12-12 Snecma Moteurs Nozzle ring with adhesive bonded blading for aircraft engine compressor
US7553130B1 (en) 2003-06-30 2009-06-30 Snecma Nozzle ring adhesive bonded blading for aircraft engine compressor
US20080022692A1 (en) * 2006-07-27 2008-01-31 United Technologies Corporation Embedded mount for mid-turbine frame
US7594404B2 (en) * 2006-07-27 2009-09-29 United Technologies Corporation Embedded mount for mid-turbine frame
DE102009007999A1 (de) * 2009-02-07 2010-08-12 Hobis Ag Leitringelement für Turbinen und Verfahren zu dessen Herstellung
WO2010094277A3 (de) * 2009-02-23 2011-06-23 Mtu Aero Engines Gmbh Gasturbinenmaschine mit einem gedämpften schaufelcluster
US20110214433A1 (en) * 2010-03-08 2011-09-08 United Technologies Corporation Strain tolerant bound structure for a gas turbine engine
US8776533B2 (en) * 2010-03-08 2014-07-15 United Technologies Corporation Strain tolerant bound structure for a gas turbine engine
US20110243742A1 (en) * 2010-03-30 2011-10-06 Snecma Stator stage for turbomachine compressor
US8794908B2 (en) * 2010-03-30 2014-08-05 Snecma Stator stage for turbomachine compressor
US9376935B2 (en) * 2012-12-18 2016-06-28 Pratt & Whitney Canada Corp. Gas turbine engine mounting ring
US20140165533A1 (en) * 2012-12-18 2014-06-19 Pratt & Whitney Canada Corp. Gas turbine engine mounting ring
US20150086352A1 (en) * 2013-09-25 2015-03-26 Pratt & Whitney Canada Corp. Gas Turbine Engine Inlet Assembly and Method of Making Same
US9784134B2 (en) * 2013-09-25 2017-10-10 Pratt & Whitney Canada Corp. Gas turbine engine inlet assembly and method of making same
US20160024971A1 (en) * 2014-07-22 2016-01-28 Rolls-Royce Plc Vane assembly
US20160025108A1 (en) * 2014-07-25 2016-01-28 Techspace Aero S.A. Vane with Sealed Lattice in a Shroud of an Axial Turbomachine Compressor
US9957980B2 (en) * 2014-07-25 2018-05-01 Safran Aero Boosters Sa Vane with sealed lattice in a shroud of an axial turbomachine compressor
RU2696177C2 (ru) * 2014-07-25 2019-07-31 Сафран Аэро Бустерс Са Осевая турбомашина
FR3097909A1 (fr) * 2019-06-27 2021-01-01 Safran Aircraft Engines Virole interne d'un carter intermédiaire, carter intermédiaire associé avec lamelles formant amortisseurs
US11352895B2 (en) 2019-10-29 2022-06-07 Raytheon Technologies Corporation System for an improved stator assembly
US11643937B2 (en) 2019-10-29 2023-05-09 Raytheon Technologies Corporation System for an improved stator assembly
US12091991B2 (en) 2019-10-29 2024-09-17 Rtx Corporation System for an improved stator assembly

Also Published As

Publication number Publication date
DE2058586A1 (de) 1971-06-24
GB1298868A (en) 1972-12-06
DE2058586B2 (de) 1976-01-15
FR2073239A1 (enrdf_load_stackoverflow) 1971-10-01

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