EP3170986B1 - Groupe d'aubes avec un dispositif de retenue circonférentiel - Google Patents

Groupe d'aubes avec un dispositif de retenue circonférentiel Download PDF

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Publication number
EP3170986B1
EP3170986B1 EP16195409.4A EP16195409A EP3170986B1 EP 3170986 B1 EP3170986 B1 EP 3170986B1 EP 16195409 A EP16195409 A EP 16195409A EP 3170986 B1 EP3170986 B1 EP 3170986B1
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EP
European Patent Office
Prior art keywords
securing
rib
guide vane
ribs
circumferential direction
Prior art date
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EP16195409.4A
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German (de)
English (en)
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EP3170986A1 (fr
Inventor
Markus Schlemmer
Oliver Thiele
Bernd Kislinger
Wilfried SCHÜTTE
Manuel Hein
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MTU Aero Engines AG
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MTU Aero Engines AG
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Publication date
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Publication of EP3170986A1 publication Critical patent/EP3170986A1/fr
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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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • 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
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2240/20—Rotors
    • F05D2240/24—Rotors for turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position

Definitions

  • the invention relates to a stator blade segment for a gas turbine, in particular an aircraft gas turbine, comprising at least one radially outer shroud and a radially inner shroud extending along a respective arc and together form a ring portion, wherein in the radial direction between the outer shroud and the inner shroud more Guide vanes are arranged side by side in the circumferential direction, which are materially connected to the inner shroud and the outer shroud, in particular integrally connected, wherein the outer shroud comprises an axial front longitudinal wall member and an axially rear end wall element, such that the outer shroud and the two end walls in longitudinal section form a trough-like profile, wherein an associated reinforcing rib is formed on the outer shroud for each vane, which extends between the two end walls.
  • the object of the invention is to improve a guide blade segment with regard to its installation and securing in an associated housing so that the above disadvantages can be overcome.
  • At least one guide vane on the rear end wall at least two adjacently arranged securing ribs are formed, wherein a limited by the two securing ribs in the circumferential direction space is tapered from radially outside to radially inside.
  • the two securing ribs are formed by a first securing rib and a second securing rib, wherein the first securing rib is connected to the reinforcing rib of the circumferentially inwardly associated associated guide vane.
  • the second securing rib is formed only in the rear end wall.
  • the configuration of the two securing ribs represents an optimized adaptation of the structural rigidity to the relevant guide blade in combination with the desired circumferential securing by the two securing ribs.
  • first securing rib and the second securing rib have mutually different rib widths relative to a width direction running along a circumferential direction tangent, wherein the circumferential direction tangent lies at the same radial distance from the center of the circle segment.
  • first securing rib and / or the second securing rib have a width that increases from radially outside to radially inside.
  • first and second securing ribs have mutually different rib heights measured at the same radial distance from the center of the circle segment in the axial direction.
  • the first securing rib at the same radial distance from the center of the circle segment have a greater rib height and a larger rib width than the second securing rib.
  • the shaping can in particular also be optimized to that effect that an improved surface pressure between the securing ribs and an adjacent housing component, in particular a housing groove receiving the two securing ribs, can be achieved.
  • the guide blade segments used in a gas turbine in particular an aircraft gas turbine, to be provided with an optimum circumferential securing for each operating state of the gas turbine.
  • the vane segment preferably comprises at least three, more preferably four to six vanes, wherein the two securing ribs are associated, relative to the circumferential direction, of an inner vane (16, 16a), preferably the second or the third or the fourth vane.
  • the vane segment may also have a different number of vanes, in particular 7 or more. More generally, the two securing ribs are arranged in a central region of the rear end wall with respect to the circumferential direction or associated with a vane adjacent to the center of the end wall or associated with an odd number of vanes of the middle vane ,
  • the two securing ribs are formed integrally with the rear end wall, in particular in one piece with the vane segment.
  • the invention also relates to a gas turbine, in particular aircraft gas turbine with at least one annular vane module, which is composed of a plurality of vane segments described above.
  • the guide vane module may be part of a turbine stage, in particular a turbine stage of a low-pressure turbine.
  • a housing element of the turbine in particular the low-pressure turbine is designed such that it is in positive or / and frictional connection with the securing ribs of at least one vane segment, such that the vane segment during operation of the gas turbine, at least in the circumferential direction is held by the connection between the housing and securing ribs. It is preferred that the connection between the housing and securing ribs is formed by a groove receiving the securing ribs.
  • FIG. 1 in simplified perspective illustrated vane element 10 comprises a radially inner shroud 12 (bottom in FIG. 1 ), a radially outer shroud 14 (at the top in FIG FIG. 1 ) and a plurality of guide vanes 16, which are arranged in the radial direction RR between the two shrouds 12 and 14. In the circumferential direction a plurality of guide vanes 16 are arranged side by side.
  • the two shrouds 14, 16 form a ring portion, wherein a plurality of Leitschaufelsegmente, which are assembled to a Leitschaufelring (not shown) in the radial direction RR and in the circumferential direction UR define an annular channel through which a fluid, in particular hot gas can flow in the axial direction AR ,
  • the guide vanes 16 are preferably materially connected to the two shrouds 12 and 14, in particular integrally formed.
  • a vane segment 10 can in particular be produced by casting from metal.
  • the guide vanes 16 are preferably hollow.
  • openings 18 can be seen, which are in communication with the cavity of the individual vanes 16 and which serve in particular to remove the casting core after the casting of the vane segment 10 from the individual vanes 16 again.
  • a front end wall 20 and a rear end wall 22 are provided in the axial direction AR, which protrude radially from the shroud 16 outwardly, such that the shroud 16 and the end walls 20, 22 a trough-like profile in a to the axial direction AR parallel longitudinal section have.
  • the end walls 20, 22 are inclined relative to the radial direction, preferably at an angle of about 20 ° to 45 °.
  • the guide vanes 16 have a flow or blade profile with a non-visible due to the viewing angle convex suction side and a concave pressure side 24, wherein the suction side and the pressure side 24 via a front edge 26 and a trailing edge 28 are interconnected.
  • Hot gas flows in a substantially axial direction AR in the flow channels 30 formed by the shrouds 12, 14 and the guide vanes 16 due to the flow profiles of the vanes 16 acts in the embodiment shown in the circumferential direction UR to the left (or counterclockwise) force acting on the Guide vane segment 10.
  • each guide vane 16 in the radially outer shroud 14 may be assigned a reinforcing rib 32 in order to support the forces acting on the shroud 14 and the end walls 20, 22.
  • the forces acting on hot gas in the circumferential direction UR forces are further supported by at least two securing ribs 34, 36 on a housing (not shown) receiving the guide vane segment 10 so that the vane segment 10 or a Leitschaufelsegmenten formed of a turbine stage of a Gastrubine in Circumference is secured.
  • first locking rib 34 and a second locking rib 36 will be described below with reference to the enlarged view of FIG. 2 explained, the dashed bordered area II of Fig. 1 equivalent.
  • the first securing rib 34 extends in the radial direction RR from an upper edge 38 of the end wall 22 downwardly or radially inwardly. In its upper region, starting from the upper edge 38, it has a transition region 35, which is preferably inclined or step-like. In its lower region, it goes directly into the vane 16a (transition region) 39 (FIG. Fig. 1 ) associated with reinforcing rib 32a.
  • the first securing rib 34 has a width extending in the circumferential direction UR or along a circumferential direction tangent, the width increasing from radially outward to radially inward. In the axial direction, the first securing rib 34 projects from the end wall 22 and has an associated height which extends in the axial direction.
  • the second securing rib 36 also extends in the radial direction RR from the upper edge 38 of the end wall 22 downwards or radially inward. In its upper region, starting from the upper edge 38, it has a transition region 37, which is preferably inclined or step-like. However, the second securing rib ends in the radial direction RR in a closing region 41 between end wall 22 and shroud 14, which is in this Presentation is only hinted at.
  • the second securing rib 36 is thus preferably provided only on the end wall 22 and has no rib-like continuation or connection to another reinforcing rib of a guide vane.
  • the second securing rib 36 also has a width extending in the circumferential direction UR or along a circumferential direction tangent, the width increasing from radially outward to radially inward. In the axial direction, the second securing rib 36 projects from the end wall 22 and has an associated height extending in the axial direction.
  • the first securing rib 34 and the second securing rib 36 are arranged at a distance RA from one another, which corresponds to a width of an intermediate space 40 formed between the two securing ribs 34, 36.
  • the width RA of the intermediate space 40 decreases from radially outside to radially inside.
  • the gap 40 is thus tapered from radially outward to radially inward or narrowing.
  • This tapering of the intermediate space 40 is formed by a first inner wall 42 of the first securing rib 34 facing the intermediate space 40 and a second inner wall 44 of the second securing rib 36 facing the intermediate space 40 being inclined relative to one another.
  • both inner walls 42, 44 are inclined, at least in relation to a plane spanned by the radial direction RR and the axial direction, which in the present representation runs essentially orthogonal to the plane of the drawing.
  • Each securing rib 34, 36 has an outer wall 46 or 48 facing away from the intermediate space 40, wherein the outer wall 46 is assigned to the first securing rib 34 and the outer wall 48 is assigned to the second securing rib.
  • the two securing ribs 34, 36 are received in a common groove formed on a housing, such that the two outer walls 46, 48 can come into contact with corresponding inner sides of the housing groove, not shown. This planar contact of the outer walls 46, 48 on the insides of the housing groove allows a support of the stator blade segment in the circumferential direction of the housing.
  • the outer walls 46, 48 of the first securing rib 34 and the second securing rib 36 preferably extend substantially parallel or slightly convergent from one another radially inward to radially outward. If the outer walls 46, 48 formed in this manner, the vane member 10 in the radial direction easily and easily be inserted into the securing ribs 34, 36 receiving groove of the housing.
  • the groove of the housing can be made particularly easy if the groove in the circumferential direction bounding walls also substantially parallel or from radially inward to radially outwardly slightly convergent to each other.
  • the width of the first securing rib 34 is greater between the radially outer transition region 35 and the radially inner transition region 39 over the entire radial length than the width of the second securing rib 36 between its radially outer transition region 37 and its termination region 41.
  • the height, d. H. the extent in the axial direction, the first securing rib 34 between the transition region 35 and the transition region 39 is greater than the height of the second securing rib 36 between the transition region 37 and the termination region 41.
  • the cross-sectional area of the first securing rib 34th This is true for the embodiment shown here for the entire radial length of the first and second securing ribs 34, 36. Accordingly, the outer surface 46 of the first securing rib 34 is larger than the outer surface 48 of the second securing rib 36 ,
  • the dimensioning of the two securing ribs takes place in consideration of the arrangement of the guide vanes 16 and their blade profile and the associated force effects in the circumferential direction on the vane segment 10.
  • act in the present embodiment when flowing through hot gas through the vane segment 10 and by a closed Leitschaufelring larger compressive forces in the circumferential direction to the left (counterclockwise), so that larger forces in the circumferential direction by means of the first securing rib 34 must be supported during operation.
  • the larger outer surface 46 of the first securing rib 34 allows sufficient surface pressure and support of the guide vane segment 10 on the housing or the groove provided in the housing.
  • the radially inwardly to radially inwardly decreasing width RA of the gap 40 is effected by the increasing width of the first and second securing ribs 34, 36.
  • the two securing ribs 34, 36 thus have their greatest width radially inward, based on the radial length of the two securing ribs 34, 36 below (radially inward) of their respective center. In this way, in particular in the transition region between the reinforcing rib 32a of the guide vane 16a acting forces that are greater than those that still act radially outward, optimally over the shroud 14, the end wall 22 and the securing ribs 34, 36 are supported.
  • the embodiment shown here is purely exemplary.
  • the first and the second securing ribs 34, 36 could also be interchanged, for example, if the guide vanes 16 were configured differently with their blade profile, in particular the pressure side and the suction side would be reversed, so that larger compressive forces would act in the circumferential direction to the right (clockwise).
  • the dimensioning of the two securing ribs can be adapted to different gas turbines or different housing. Different shapes and dimensions of the two securing ribs 34, 36 can already be taken into account during the production of a guide blade segment 10, so that a finished, in particular cast guide blade element 10 already has the two securing ribs 34, 36, which are formed integrally with the guide blade segment.
  • an optimized setting of the surface pressure between the securing ribs and the housing groove can be achieved by the choice and dimensioning of the two securing ribs. Furthermore, there is a cost savings in the housing production by choosing an optimized distance between the two securing ribs, by the elimination of additional processing steps, such as application of solder material or insertion of locking pins and the like.
  • the two securing ribs 34, 36 thus allow a total of flexible adaptation to the necessary structural conditions of a gas turbine type, in which the guide blade segments are to be used.

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

Claims (12)

  1. Segment d'aubes directrices (10) destiné à une turbine à gaz, en particulier une turbine à gaz d'avion, comprenant au moins une bande de recouvrement (14) radialement extérieure et une bande de recouvrement (12) radialement intérieure qui s'étendent le long d'un arc de cercle respectif et forment conjointement une portion annulaire, plusieurs aubes directrices (16, 16a) étant disposées, dans une direction radiale (RR), les unes à côté des autres dans la direction circonférentielle (UR) entre la bande de recouvrement extérieure (14) et la bande de recouvrement intérieure (12) et étant raccordées par liaison de matière à la bande de recouvrement intérieure (12) et à la bande de recouvrement extérieure (14), en particulier d'une seule pièce, la bande de recouvrement extérieure (14) comprenant dans une direction longitudinale axiale (AR) un élément de paroi frontale (20) axialement antérieur et un élément de paroi frontale (22) axialement postérieur de telle sorte que la bande de recouvrement extérieure (14) et les deux parois frontales (20, 22) forment en coupe longitudinale un profil en forme d'auge, une nervure de renforcement (32, 32a) associée étant formée au niveau de la bande de recouvrement extérieure (14) pour chaque aube directrices (16, 16a) et s'étendant entre les deux parois frontales (20, 22), au moins deux nervures de blocage (34, 36) juxtaposées étant constituées, pour au moins une aube directrices (16a), au niveau de la paroi frontale postérieure (22), un espace intermédiaire (40) étant délimité dans la direction circonférentielle (UR) par les deux nervures de blocage (34, 36), caractérisé en ce qu'une largeur (RA) de l'espace intermédiaire (40) diminue depuis le côté radialement extérieur vers le côté radialement intérieur de telle sorte que l'espace intermédiaire (40) soit formé de manière à se rétrécir depuis le côté radialement extérieur vers le côté radialement intérieur.
  2. Segment d'aubes directrices selon la revendication 1, caractérisé en ce que les deux nervures de blocage (34, 36) sont formées par une première nervure de blocage (34) et une deuxième nervure de blocage (36), la première nervure de blocage (34) étant raccordée à une nervure de renforcement (32a) qui est associée à une aube directrices (16a) située intérieurement dans la direction circonférentielle (UR).
  3. Segment d'aubes directrices selon la revendication 2, caractérisé en ce que la deuxième nervure de blocage (36) est formée seulement dans la paroi frontale postérieure (22).
  4. Segment d'aubes directrices selon la revendication 2 ou 3, caractérisé en ce que la première nervure de blocage (34) et la deuxième nervure de blocage (36) ont des largeurs différentes par rapport à une direction de largeur s'étendant le long d'une tangente à la direction circonférentielle, la tangente à la direction circonférentielle se situant respectivement à la même distance radiale du centre du segment de cercle.
  5. Segment d'aubes directrices selon la revendication 4, caractérisé en ce que la première nervure de blocage (34) et/ou la deuxième nervure de blocage (36) ont une largeur qui augmente depuis le côté radialement extérieur vers le côté radialement intérieur.
  6. Segment d'aubes directrices selon l'une des revendications 2 à 5, caractérisé en ce que la première et la deuxième nervure de blocage (34, 36) ont des hauteurs différentes mesurées à la même distance radiale du centre du segment de cercle dans la direction axiale (AR).
  7. Segment d'aubes directrices selon la revendication 6, caractérisé en ce que la première nervure de blocage (34) a, à la même distance radiale du centre du segment de cercle, une hauteur et une largeur qui sont supérieures à celles de la deuxième nervure de blocage (36).
  8. Segment d'aubes directrices selon l'une des revendications précédentes, caractérisé en ce qu'il comprend au moins trois, de préférence quatre à six, aubes directrices (16, 16a), les deux nervures de blocage (34, 36) étant associées à une aube directrice (16, 16a) située intérieurement, de préférence à la deuxième ou à la troisième ou à la quatrième aube directrice (16a), par rapport à la direction circonférentielle (UR).
  9. Segment d'aubes directrices selon l'une des revendications précédentes, caractérisé en ce que les deux nervures de blocage (34, 36) sont formées d'une seule pièce avec la paroi frontale postérieure (22), en particulier d'une seule pièce avec le segment d'aube directrice(10).
  10. Turbine à gaz, en particulier turbine à gaz d'avion, comprenant au moins un module d'aube directrice de forme annulaire qui se compose de plusieurs segments d'aube directrice (10) selon l'une des revendications précédentes.
  11. Turbine à gaz selon la revendication 10, caractérisée en ce que le module d'aube directrice fait partie d'un étage de turbine, en particulier d'un étage de turbine d'une turbine à basse pression.
  12. Turbine à gaz selon la revendication 10 ou 11, caractérisée en ce qu'un élément de carter de la turbine, en particulier de la turbine à basse pression, est formée de manière à être raccordé par complémentarité de formes et/ou liaison de friction aux nervures de blocage (34, 36) d'au moins un segment d'aube directrice (10) de telle sorte que, pendant le fonctionnement de la turbine à gaz, le segment d'aube directrice (10) est retenu au moins dans la direction circonférentielle par le raccordement entre le carter et les nervures de blocage (34, 36), le raccordement entre le carter et les nervures de blocage (34, 36) étant de préférence formée par une rainure recevant les nervures de blocage (34, 36) .
EP16195409.4A 2015-11-19 2016-10-25 Groupe d'aubes avec un dispositif de retenue circonférentiel Active EP3170986B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102015222834.7A DE102015222834A1 (de) 2015-11-19 2015-11-19 Schaufelcluster mit Umfangssicherung

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EP3170986A1 EP3170986A1 (fr) 2017-05-24
EP3170986B1 true EP3170986B1 (fr) 2018-12-12

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US (1) US10428668B2 (fr)
EP (1) EP3170986B1 (fr)
DE (1) DE102015222834A1 (fr)
ES (1) ES2704288T3 (fr)

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KR101901683B1 (ko) * 2017-02-06 2018-09-27 두산중공업 주식회사 직선형 냉각홀을 포함하는 가스터빈 링세그먼트 및 이를 포함하는 가스터빈
US11536147B2 (en) 2021-03-30 2022-12-27 Raytheon Technologies Corporation Vane arc segment with flange and gusset

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US3403889A (en) * 1966-04-07 1968-10-01 Gen Electric Frame assembly having low thermal stresses
EP1112440B1 (fr) * 1998-08-31 2003-06-18 Siemens Aktiengesellschaft Pale directrice de turbine
DE10331599A1 (de) * 2003-07-11 2005-02-03 Mtu Aero Engines Gmbh Bauteil für eine Gasturbine sowie Verfahren zur Herstellung desselben
CA2633337C (fr) * 2005-12-29 2014-11-18 Rolls-Royce Power Engineering Plc Profil d'une aube de guidage de tuyere de second etage
FR2928962B1 (fr) * 2008-03-19 2013-10-18 Snecma Distributeur de turbine a pales creuses.
FR2941488B1 (fr) * 2009-01-28 2011-09-16 Snecma Anneau de turbine a encoche anti-rotation
US8360716B2 (en) 2010-03-23 2013-01-29 United Technologies Corporation Nozzle segment with reduced weight flange
PL2615243T3 (pl) * 2012-01-11 2017-12-29 MTU Aero Engines AG Segment wieńca łopatkowego do maszyny przepływowej i sposób jego wytwarzania
FR2990719B1 (fr) * 2012-05-16 2016-07-22 Snecma Distributeur de turbomachine, et procede de fabrication
EP2787178B1 (fr) 2013-04-03 2016-03-02 MTU Aero Engines AG Ensemble d'aube directrice
US9797262B2 (en) 2013-07-26 2017-10-24 United Technologies Corporation Split damped outer shroud for gas turbine engine stator arrays

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US10428668B2 (en) 2019-10-01
DE102015222834A1 (de) 2017-05-24
ES2704288T3 (es) 2019-03-15
US20170145842A1 (en) 2017-05-25
EP3170986A1 (fr) 2017-05-24

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