EP2261462A1 - Turbinenstufenendwand - Google Patents

Turbinenstufenendwand Download PDF

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Publication number
EP2261462A1
EP2261462A1 EP09161706A EP09161706A EP2261462A1 EP 2261462 A1 EP2261462 A1 EP 2261462A1 EP 09161706 A EP09161706 A EP 09161706A EP 09161706 A EP09161706 A EP 09161706A EP 2261462 A1 EP2261462 A1 EP 2261462A1
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EP
European Patent Office
Prior art keywords
channel
turbine stage
adjacent
flow
face
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.)
Withdrawn
Application number
EP09161706A
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English (en)
French (fr)
Inventor
Benjamin Megerle
Thomas Mokulys
Said Havakechian
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.)
General Electric Technology GmbH
Original Assignee
Alstom Technology AG
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP09161706A priority Critical patent/EP2261462A1/de
Priority to JP2010125872A priority patent/JP2010281320A/ja
Priority to CN201010233595.7A priority patent/CN101922311A/zh
Priority to US12/792,352 priority patent/US20100303627A1/en
Publication of EP2261462A1 publication Critical patent/EP2261462A1/de
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • 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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • 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
    • F05D2250/60—Structure; Surface texture
    • F05D2250/61—Structure; Surface texture corrugated
    • F05D2250/611—Structure; Surface texture corrugated undulated
    • 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
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved

Definitions

  • the disclosure relates generally to axial gas and steam turbines in which there are one or more rows of generally radially extending airfoils of non rotating vanes and rotating blades. More specifically the disclosure relates to configurations of end walls joined to radial ends of airfoils with improved aerodynamic behaviour.
  • end wall is broadly defined as any surface at a radial end of an airfoil and from which the airfoil radially extends. End walls thus include but are not limited to airfoil platforms and shrouds.
  • the ideal flow through a turbine is termed the "primary flow” wherein the difference between the primary flow and the actual flow is termed the “secondary flow”.
  • the secondary flow represents, to a large extent, a loss that has a major impact on axial turbine efficiency.
  • the pressure side leg of the vortex is influenced by the airfoil-to-airfoil pressure gradient as it enters the flow passage and travels towards the suction side.
  • the resulting cross-passage flow along the end wall imposes vortex motion in the cascade.
  • These vortices may commonly be referred to as passage vortices that include horseshoe vortices in their core.
  • These vortices can be present in any flow channel with a curved shape and a boundary layer.
  • the strength of this secondary flow in a cascade is dependent on a number of other factors including the amount of turning and the shape of the incoming boundary layer.
  • the end wall vortex generation negatively affects turbine efficiency, contributing up to 35% of the total losses for a typical high-pressure turbine.
  • the key cause for the additional loss generation is the passage vortex that grows downstream of the cascade.
  • the kinetic energy stored in this vortex is lost for further use as it is mostly mixed out downstream.
  • the passage vortex can be easily detected as a high loss core existing away from the suction surface close to the centre of the passage vortex.
  • Airfoil design has evolved to reduce secondary flow by optimising the three-dimensional shape of airfoils and more recently by contouring of the end walls.
  • This technology referred to as Tangential End Wall Contouring (TEWC)
  • TEWC Tangential End Wall Contouring
  • US patent application US 2007/0059177 A1 describes such an end wall nonaxisymmetric profile. This solution comprises forming circumferentially extending sinusoids at a number of axial positions wherein corresponding points on successive sinusoids are joined by spline curves so that the curvature of the end wall is smooth.
  • Another alternative solution is to provide the end wall with a fence that lifts the vortex up off the end wall and into the main flow, which has the effect of washing out the vortices.
  • the fence of which an example is described in ASME Turbo Expo 2000 " Secondary flow measurements in a turbine passage with end wall flow modification " 2000-GT-0212 has a leading edge at the centre of a line connecting the leading edges of the pressure and suction side airfoils. While such walls can reduce aerodynamic losses practical problems can arise due to the need to cool the fence.
  • EP 1 995 410 A1 provides a solution in which an end wall of a turbine stage cascade includes a first projection having a ridge extending downward from the trailing edge of a turbine blade toward the downstream side, gently at the beginning and steeply at the end, and along the suction side of an adjacent turbine blade. Such an arrangement is however limited by the fact that downstream axial space is required and therefore such a solution may not always be applicable.
  • the disclosure is directed towards the problem, in a turbine, of over and under turning capability and/or reduced helicity losses as a result of secondary flows caused by cross flow that flows in the pitchwise direction from a pressure face of an airfoil towards the suction face of an adjacent airfoil across an end wall surface.
  • An aspect of the invention is based on the principle of one or more adjacent channels formed in end walls in flow passages between adjacent airfoils.
  • Each channel extends in the primary flow direction and may be preferably located adjacent airfoil pressure faces.
  • the channels each have two angled walls, which in conjunction with the configuration and location of the channels, are configured to reduce the potential for secondary flow formation in the channels.
  • a turbine stage comprising a circumferentially distributed row of adjacent airfoils each with a pressure face, a suction face and, one end wall, from which the airfoils radially extends or two end walls between which the airfoils extend.
  • the turbine stage further has a flow passage defined by a region between a pressure face of a first airfoil, a suction face of an adjacent second airfoil, a leading edge line, define as a line extending between the lead edges of adjacent airfoils, and a trailing edge line, defined as a line extending between the trailing edges of adjacent airfoils.
  • the flow passage has a surface, which in its unmodified form defines a datum.
  • the turbine stage, in each flow passage has one or, two or more adjacent channels, adjacent a pressure face, that modify the surface and extend in the direction of primary flow lines from a point towards the leading edge line to a point towards the trailing edge line.
  • Each channel consists of two channels walls angled relative to the datum that provide the channels with have a low point, two high points, and a channel height which is the radial distance between the low point and highest of the high points.
  • the location of the channels adjacent the pressure face reduces the extent of influence of cross flow pitchwise across the flow passage so by reducing the influence of secondary flow. The closer the channel is to the pressure face the more pronounced this effect. In this way any negative effect on aerodynamic performance is more than offset by the benefit of reduced secondary cross flow.
  • the high points of each channel do not extend above the datum, reducing the impact of the channels on primary flow thus reducing scraping losses.
  • the low point of each channel is substantially in the midpoint, pitchwise, between high points of each channel while in another aspect, the angle of the walls of each channel closer to the pressure face is less, relative to the datum, than channel walls closer to the suction face.
  • the channel height in the primary flow direction, increases to a maximum at a relative channel length, in the direction of primary flow lines, of between 0.35-0.55 at which point the channel height decreases. In the last fifth of the length of the channel this rate of decrease may be less. In this way the channel depth provides a balance between scraping losses, which change with the velocity profile in the flow passage, and cross flow presence.
  • each successive adjacent channel adjacent in the pitch wise direction extending from the pressure face to the suction face, remains the same or decreases.
  • the channel height of the channel adjacent the pressure face is at least twice that of the channel furthest, in the pitch wise direction, from the pressure face. In this way channels are configured to reduce cross flow where its affect is strongest i.e. adjacent the pressure face.
  • each flow passage in each flow passage, the point of extension of each adjacent channel is the same or further from the leading edge line the closer the channel is to the suction face thus defining an extended region, adjacent the suction face, which is free of channels.
  • the extended region can be generally defined as a region, that in operation, is configured in a region that is essentially free of secondary flow vortices caused by cross flow originating from the pressure face.
  • FIGs. 1 and 2 show top and perspective views respectively of two adjacent airfoils 10 of a turbine stage, in which airfoils 10 are adjacently and circumferentially distributed in rows.
  • Each airfoil 10 is integrally joined at one or both radial ends to corresponding end walls 12, partially shown as grid lines.
  • the area between the pressure face 14 and suction face 16 of adjacent airfoils 10 defines a flow passage 18 further bound by a region extending between a leading edge line 20, define as a line extending between the lead edges 21 of adjacent airfoils 10, and a trailing edge line 22, defined as a line extending between the trailing edge 23 of adjacent airfoils 10.
  • the flow passage 18 has a surface common with a surface of the end walls 12.
  • the surface In its unmodified form the surface defines a datum DR, shown in FIG. 4 , wherein "unmodified form” means the contour the passage surface would take if the surfaces were not changed, for example, by TEWCs.
  • the grid lines in FIGs. 1 and 2 representing an unmodified surface comprise of, primary flow lines PFL that represent ideal lines of flow unaffected by secondary flow, and, pitch wise sections A-D.
  • FIG. 3 shows an exemplary embodiment applied to the turbine stage shown in FIGs. 1 and 2 . It comprises channels 30 formed in the passage surfaces of end walls 12, at one or both radial end of adjacent airfoils 10.
  • the channels 30 extend in the direction of primary flow lines PFL and thus are substantially parallel to each other. The extension is from a point towards the leading edge line 20 to a point towards the trailing edge line 22.
  • Each channel 30 consists of two channel walls 32 that angle relative to the datum DR, shown in more details in FIGs 5 and 6 , and join to define a low point LP of the channel 30 relative to the datum DR.
  • FIG. 4 shows an exemplary cross section through a pitch wise section A-D extending between the pressure face 14 of one airfoil 10 to the suction face 16 of another adjacent airfoil 10.
  • Each channel 30 has two channel walls 32, one closer to the pressure face 14 and the other closer to the suction face 16.
  • the channel height CH is the radial height, that is the height measured perpendicular to the datum DR, between the channel's low points LP and the highest high point HP.
  • “low” and " high” are relative to the datum wherein “low” refers to a negative extension from the datum DR into the end wall 12, while “high” refers to a positive extension in the direction away from the end wall 12.
  • the indication is independent of absolute location. That is, even though the high point HP extends in a direction away from the end wall 12, the high point HP, as shown in FIG. 7 , may or may not extend above the height of the datum DR
  • the "channel wall angle” e is the angle of a nominal channel wall 33 relative to the datum DR wherein the nominal channel wall 33, without curvature, approximates the actual channel wall 32.
  • FIG. 5 showing an expanded view of V of FIG. 4
  • the wall angle e is taken to be the angle of the nominal channel wall 33, which is the average angle of the nominal channel wall 32.
  • FIG. 6 which is an expanded view of VI of FIG. 4
  • a channel wall 32 is shown with a rounded-off end section that is otherwise straight. In this case the nominal channel wall 33 corresponds to the channel wall 32 straight portion, disregarding the rounded end section.
  • the purpose of the channels 30 is to reduce cross flow and so reduce secondary flow and resulting losses.
  • the preferred channel height CH and preferred number of channels 30 is dependent on the degree of cross flow, estimatable using known techniques, described, for example in Harvey, N. W. et al, 2000 " Nonaxisymmetric Turbine end Wall Design: Part I "ASME J. Turbomach., 122, pp. 278-285 and, Hartland, J. C. et al, 2000 " Nonaxisymmetric Turbine End wall Design: Part II "ASME 122 J. Turbomach, 122, pp. 286-293 .
  • With increasing channel height CH and channel 32 number passage surface area increases, which, in the absences of secondary flow, results in increased scraping losses.
  • An embodiment in its simplest form suitable for turbine stages with minimal cross flow therefore comprises one channel 30 located adjacent the pressure surface 14, which is the region with the most significant cross flow.
  • the channel depth CH is a function of the number of channels 30 and the degree of cross flow. If the channel depth CH is too great further secondary flow can be created resulting in additional losses. If the channel depth CH is too low the ability of the channel to limit cross flow will be limited. A further consideration is channel wall angle ⁇ . If too steep, additional secondary flow may be created. Channel design is therefore a compromise between at least these factors and so is strongly dependent on airfoil design and operation conditions. In consideration of these factors an optimum design can be derived by simulation using known methods.
  • the low point LP of each channel 30 is at the pitchwise midpoint between the high points HP of the channel as shown in FIG. 4 .
  • the channel height CH of each successive adjacent channel 30 in the pitchwise direction from the pressure face 14 to the suction face 16 remains the same or decreases.
  • the channel height CH of the channel 30 adjacent the pressure face 14 is at least twice that of the channel 30 furthest from the pressure face 14, as can be seen in FIG. 4 . As cross flow is typically greatest towards the pressure face 14 the benefit of channels 30 decreases towards the suction face 16.
  • the low point is closer to the suction face 16, represented as the pitchwise position "1" in FIG. 7 than the pressure face 14, which is represented as "0".
  • This typically results in the channel wall angle e of the channel wall 32 closer the suction face 16 being greater than the channel wall angle e of the channel wall 32 closer the pressure face 14.
  • the channel wall angle e of the channel wall 32 located closer to the suction face 16 is typically less than less than 90 degrees as an angle approaching 90 degrees or greater may create additional vortices resulting in additional losses.
  • the channel walls 32 are configured such that the channels 20 do not extend above the datum DR, as shown in FIG. 7 wherein "0" is the channel height CH at the datum DR.
  • the flow is accelerated significantly.
  • Scraping losses which have a squared relationship with velocity, are of greatest significance in the region of highest velocity.
  • the highest velocity may correspond to a region where the separation distance measured in the pitchwise direction, between adjacent airfoils 10, is smallest.
  • overall efficiency may be optimised if the channel height CH is limited so as to be lower than would optimally be designed in view only of predicted cross flow. Therefore, in an exemplary embodiment, in the direction of primary flow lines PFL extending from towards the leading edge line 20 to the trailing edge line 20, the channel height initially increases to a maximum at a relative channel length of between 0.35-0.55 after which is decreases.
  • the decrease is not as pronounced in the last fifth of the relative channel length.
  • the relative channel length is the length point along a channel 30 measured relative to the total length of the channel 30.
  • FIG. 8 shows an example of one configuration of these embodiments in which it was found that for one set of operating conditions not only can scraping losses be reduced but also over and under turning performance can be slightly improved without detrimentally affecting helicity.
  • FIG. 9 shows an exemplary embodiment where the channels 30 towards the suction face 16 starts further from the leading edge line 20 than channels 30 closer to the pressure face 14. That is, their point of extension from the leading edge line 20 is further.
  • the extended region ER may be bounded by a midpoint on the leading edge line 20, a point along the suction face 16 and a point on the suction face 16 at which the suction face 16 and leading edge line 20 join, as shown in both FIGs. 9 and 10 .
  • Such an arrangement is beneficial when the flow across the extended region ER is essential free of secondary flow, as shown in FIG. 10 , and as such the loss in this region primarily comprises scraping losses.
  • the extended region ER is the region adjacent the suction face 16 towards the leading edge line 20 that is essential free of secondary flow as shown by the flow lines FL in FIG. 10 .
  • the extended region ER is defined by a region derived and determined by known flow simulation methods.
  • FIGs. 11-14 shows the performance that can be achieved with a combination of the various exemplary embodiments. Improvements include over and under turning, shown in FIGs. 11 and 12 for both a stator and a rotor, and helicity, shown in FIGs. 13 and 14 also for a stator and rotor.

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  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP09161706A 2009-06-02 2009-06-02 Turbinenstufenendwand Withdrawn EP2261462A1 (de)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP09161706A EP2261462A1 (de) 2009-06-02 2009-06-02 Turbinenstufenendwand
JP2010125872A JP2010281320A (ja) 2009-06-02 2010-06-01 タービン段
CN201010233595.7A CN101922311A (zh) 2009-06-02 2010-06-02 涡轮级
US12/792,352 US20100303627A1 (en) 2009-06-02 2010-06-02 Turbine stage

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09161706A EP2261462A1 (de) 2009-06-02 2009-06-02 Turbinenstufenendwand

Publications (1)

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EP2261462A1 true EP2261462A1 (de) 2010-12-15

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EP09161706A Withdrawn EP2261462A1 (de) 2009-06-02 2009-06-02 Turbinenstufenendwand

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US (1) US20100303627A1 (de)
EP (1) EP2261462A1 (de)
JP (1) JP2010281320A (de)
CN (1) CN101922311A (de)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2012097798A1 (de) * 2011-01-19 2012-07-26 Mtu Aero Engines Gmbh Zwischengehäuse einer gasturbine mit einer aussen liegenden begrenzungswand welches stromaufwärts einer stützrippe eine in umfangrichtung verändernde kontur aufweist zur verringerung der sekundärströmungsverluste
EP2597257A1 (de) * 2011-11-25 2013-05-29 MTU Aero Engines GmbH Beschaufelung
EP2518269A3 (de) * 2011-04-28 2013-11-27 Hitachi Ltd. Leitschaufel einer Gasturbine
EP2806102A1 (de) 2013-05-24 2014-11-26 MTU Aero Engines GmbH Schaufelgitter einer Strömungsmaschine und zugehörige Strömungsmaschine
EP2835499A1 (de) 2013-08-06 2015-02-11 MTU Aero Engines GmbH Schaufelgitter und zugehörige Strömungsmaschine
FR3014943A1 (fr) * 2013-12-18 2015-06-19 Snecma Piece de turbomachine a surface non-axisymetrique
US9194235B2 (en) 2011-11-25 2015-11-24 Mtu Aero Engines Gmbh Blading
EP3091180A1 (de) * 2015-05-08 2016-11-09 MTU Aero Engines GmbH Leitschaufelsegment mit konturierter endwand
US20180252107A1 (en) * 2017-03-03 2018-09-06 MTU Aero Engines AG Contouring a blade/vane cascade stage

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US8807930B2 (en) 2011-11-01 2014-08-19 United Technologies Corporation Non axis-symmetric stator vane endwall contour
EP2639004B1 (de) * 2012-03-15 2016-05-11 MTU Aero Engines GmbH Schaufelkranzsegment mit Ringraumbegrenzungsfläche mit einem welligen Höhenprofil sowie Verfahren zur Herstellung desselben
US9085985B2 (en) 2012-03-23 2015-07-21 General Electric Company Scalloped surface turbine stage
US9598967B2 (en) * 2012-12-18 2017-03-21 United Technologies Corporation Airfoil member and composite platform having contoured endwall
EP2746533B1 (de) * 2012-12-19 2015-04-01 MTU Aero Engines GmbH Schaufelgitter und Strömungsmaschine
EP2806103B1 (de) * 2013-05-24 2019-07-17 MTU Aero Engines AG Schaufelgitter und Strömungsmaschine
CN104005796B (zh) * 2014-05-09 2015-12-30 上海交通大学 新型涡轮叶栅端壁的沟槽减损结构和方法
KR20190046118A (ko) * 2017-10-25 2019-05-07 두산중공업 주식회사 터빈 블레이드
US11560797B2 (en) * 2018-03-30 2023-01-24 Siemens Energy Global GmbH & Co. KG Endwall contouring for a conical endwall
US10968748B2 (en) * 2019-04-08 2021-04-06 United Technologies Corporation Non-axisymmetric end wall contouring with aft mid-passage peak
US10876411B2 (en) * 2019-04-08 2020-12-29 United Technologies Corporation Non-axisymmetric end wall contouring with forward mid-passage peak
CN110145370A (zh) * 2019-04-30 2019-08-20 浙江大学 一种吸力面波浪形的低压涡轮叶片
US20210079799A1 (en) * 2019-09-12 2021-03-18 General Electric Company Nozzle assembly for turbine engine
CN111946666B (zh) * 2020-07-20 2022-04-19 中国科学院工程热物理研究所 一种轴流压气机端壁附面层流动调控结构

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Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2012097798A1 (de) * 2011-01-19 2012-07-26 Mtu Aero Engines Gmbh Zwischengehäuse einer gasturbine mit einer aussen liegenden begrenzungswand welches stromaufwärts einer stützrippe eine in umfangrichtung verändernde kontur aufweist zur verringerung der sekundärströmungsverluste
US9382806B2 (en) 2011-01-19 2016-07-05 Mtu Aero Engines Gmbh Intermediate housing of a gas turbine having an outer bounding wall having a contour that changes in the circumferential direction upstream of a supporting rib to reduce secondary flow losses
US9334745B2 (en) 2011-04-28 2016-05-10 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine stator vane
EP2518269A3 (de) * 2011-04-28 2013-11-27 Hitachi Ltd. Leitschaufel einer Gasturbine
US9963973B2 (en) 2011-11-25 2018-05-08 Mtu Aero Engines Gmbh Blading
US9194235B2 (en) 2011-11-25 2015-11-24 Mtu Aero Engines Gmbh Blading
US9316103B2 (en) 2011-11-25 2016-04-19 Mtu Aero Engines Gmbh Blading
EP2597257A1 (de) * 2011-11-25 2013-05-29 MTU Aero Engines GmbH Beschaufelung
US9745850B2 (en) 2013-05-24 2017-08-29 MTU Aero Engines AG Blade cascade and continuous-flow machine
EP2806102A1 (de) 2013-05-24 2014-11-26 MTU Aero Engines GmbH Schaufelgitter einer Strömungsmaschine und zugehörige Strömungsmaschine
EP2835499A1 (de) 2013-08-06 2015-02-11 MTU Aero Engines GmbH Schaufelgitter und zugehörige Strömungsmaschine
US10041353B2 (en) 2013-08-06 2018-08-07 MTU Aero Engines AG Blade cascade and turbomachine
FR3014943A1 (fr) * 2013-12-18 2015-06-19 Snecma Piece de turbomachine a surface non-axisymetrique
US10519980B2 (en) 2013-12-18 2019-12-31 Safran Aircraft Engines Turbomachine component or collection of components and associated turbomachine
EP3091180A1 (de) * 2015-05-08 2016-11-09 MTU Aero Engines GmbH Leitschaufelsegment mit konturierter endwand
US10598032B2 (en) 2015-05-08 2020-03-24 MTU Aero Engines AG Gas turbine guide vane element
US20180252107A1 (en) * 2017-03-03 2018-09-06 MTU Aero Engines AG Contouring a blade/vane cascade stage
US10648339B2 (en) * 2017-03-03 2020-05-12 MTU Aero Engines AG Contouring a blade/vane cascade stage

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US20100303627A1 (en) 2010-12-02
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