EP0943784A1 - Veine profilée pour une turbomachine axiale - Google Patents

Veine profilée pour une turbomachine axiale Download PDF

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Publication number
EP0943784A1
EP0943784A1 EP98810233A EP98810233A EP0943784A1 EP 0943784 A1 EP0943784 A1 EP 0943784A1 EP 98810233 A EP98810233 A EP 98810233A EP 98810233 A EP98810233 A EP 98810233A EP 0943784 A1 EP0943784 A1 EP 0943784A1
Authority
EP
European Patent Office
Prior art keywords
blade
contour
guide vane
rotor
guide
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Withdrawn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
EP98810233A
Other languages
German (de)
English (en)
Inventor
Said Dr. Havakechian
Kurt C. Dr. Heiniger
Peter Szincsak
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.)
ABB Asea Brown Boveri Ltd
ABB AB
Original Assignee
ABB Asea Brown Boveri Ltd
Asea Brown Boveri AB
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 ABB Asea Brown Boveri Ltd, Asea Brown Boveri AB filed Critical ABB Asea Brown Boveri Ltd
Priority to EP98810233A priority Critical patent/EP0943784A1/fr
Publication of EP0943784A1 publication Critical patent/EP0943784A1/fr
Withdrawn legal-status Critical Current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

Definitions

  • the invention relates to axially flow-through turbomachines with a multi-stage Blading.
  • the invention relates to contouring and design of the flowed channel in its bladed area. It is applicable for blading with tip seal and for those with cover plate or Shroud seals.
  • the main flow (the flue gas) flows in turbomachines with axial flow with high kinetic energy through the fixed guide vanes of these advantageously directed to the subsequent row of rotating blades. This gives the latter a strong impulse in the circumferential direction. over the expansion becomes part of the thermal energy of each blade row Main flow converted into mechanical energy.
  • blading and the associated redirection of the mass particles is axial flowed through turbo machines the resulting flow turbulent.
  • the reaction vanes When the reaction vanes are subject to vortex flow, they flow axially through them Turbo machines with cylindrical blades are constantly changing Circumferential component of the flow velocity meandering trajectories of the mass particles. This wavy flow is in the book "Thermal Turbomachinery" by Walter Traupel, Volume 1, Springer Verlag 1966, chapter 7 described. About gap losses between blading and diminish boundary walls, it is while maintaining this undulating Flow known, the stator-side and the rotor-side limitation with approximately the same ripple as the current has.
  • the stator-side waveform be formed in that the contour in the area of the guide vane root to the machine longitudinal axis and in the area the blade tip is directed away from the machine longitudinal axis. Accordingly the rotor-side waveform is then formed in that the contour is executed in the area of the guide blade tip to the machine longitudinal axis and directed away from the machine longitudinal axis in the area of the blade root is.
  • the invention is based, with a blading the task at the beginning mentioned type to create a channel contour, through which the efficiency at the same Blade load can be increased or the possible blade load can be increased with constant efficiency.
  • the advantages of the invention are, for example, a reduction in the diffusion flow at the vane tip through enhanced streamlined convergence, in an increased degree of load absorption in the rear area of the Guide blade root (Degree of Aft-Loading or DAL for short), a reinforced Degree of load absorption in the rear area of the blade root, reduced Diffusion flow in the area of the blade tip and reinforced Degree of load absorption in the rear area of the blade tip. Furthermore the invention leads to a reduction in leakage currents and an improved Backflow of these leakage currents into the main flow.
  • the concave contour on the blade feet is particularly useful to have the concave contour on the blade feet to combine the convex contour at the guide vane tips, as this leads to a Homogenization of the outflow from the blades with regard to total pressure and discharge angle leads. Furthermore, the concave is particularly advantageous Contour on the blade roots or the convex contour on the guide blade tips with the concave contour on the guide vane feet or the conical Combine contour at the blade tips.
  • transitions of the contours between the blade tip and blade root or Blade root and blade tip are, depending on the used Contour, preferably chosen so that it has a streamlined transition Enable working medium.
  • they are downstream or upstream end areas of the individual contours are coordinated so that the labyrinth flow passing the blade tips into the area the recessing of the concave contour of the subsequent blade root flows in. This inflow is essentially parallel to the main flow. The direction of flow of the main flow in this area becomes primarily through the appropriately trained flow restricting Wall of the previous blade tip determined.
  • the invention is both for high pressure turbines and for medium pressure turbines applicable. It can also be used in the first rows of blades of low-pressure turbines be applied.
  • Fig. 1 shows a longitudinal section through a turbine. There is only a section a multi-stage medium pressure blading shown. That flows in the drawing Working medium from left to right through channel 2.
  • a blade stage consisting of guide blade Le2 and rotor blade La2.
  • the rotor blade La1 of the upstream blade stage is shown.
  • the associated guide blade Le1 is not shown.
  • the main flow of the working medium (the flue gas) flows through the stationary with high kinetic energy Guide blades Le2 and from these to the following row Blades La2 directed. This gives the latter a strong impact Impulse in the circumferential direction.
  • the expansion becomes a part of each row of blades the thermal energy of the main flow is converted into mechanical energy.
  • the energy absorbed by the blades is ultimately transferred to the Transfer rotor.
  • the blade roots 22 of the blades 12 have their radially inner side End one base 30 each.
  • the base 30 are in recesses 34 of the Rotors 4 used.
  • At the radially outer end of the blades 12 is the blade tips 20.
  • the blade tips 20 of an impeller are connected to each other by means of only indicated cover plates.
  • the radial outer contours 38 of the cover plates can be geometrical depending on the row of runs be designed differently. They seal with their contours to form Labyrinths 42 against the sealing strips 40, which are in the stator 6 in a known manner are arranged.
  • Partial flows of the working medium the so-called leakage currents, occur after flowing out of a guide vane via the labyrinth inlet 44 the labyrinth 42 and from there they flow again via the labyrinth outlet 46 back into the working medium.
  • the occurrence of leakage currents is high Losses connected. On the one hand they get lost in the working medium, on the other hand they cause 2 eddies in the working medium when they flow back into the channel. This is with secondary losses in the downstream blade stage connected.
  • a guide vane row is constructed analogously.
  • the guide vane feet 18 of the Guide vanes 14 each have a base 32 at their radially outer end on.
  • the bases 32 are inserted in recesses 36 of the stator 6.
  • the guide vane tips are located on the inside end of the guide vanes 14 16.
  • the guide vane tips 16 of a row of guide vanes are only indicated by means of Cover plates connected together.
  • the radially inner contours 48 of the cover plates can have different geometries depending on the row of runs be. They seal with their contours to form labyrinths 52 the sealing strips 50, which are arranged in the rotor 4 in a known manner. Also here leakage currents enter the labyrinth 52 via the labyrinth inlet 54 and from there they flow back into the working medium via the labyrinth outlet 56 back.
  • the geometric shape of the labyrinths, as well as the labyrinth entries and exits determined by the blade feet and blade tips of the directly adjacent ones Blading and by stator and rotor parts, which lead the current in the take over areas not bladed.
  • the main flow of the working medium flows through the channel 2. This has a conical outer in Fig. 1 Baseline on the stator 6 and a cylindrical inner baseline on the rotor 4. However, neither is mandatory.
  • stator-side flow-limiting wall 10 becomes walls of the flowed channel 2 in the area of the rotor blade through the Channel facing side of the cover plate of the blade tips 20 and in the area of the guide vane blade through the side facing the duct that is not in detail shown base plates of the vane feet 18 formed. Accordingly, the rotor-side flow-limiting wall 8 of the channel 2 through which flow flows in the region of the rotor blade blade through the side facing the duct, which is not in the Detail shown base plates of the blade roots 22 and in the area of the guide blade blade through the side of the cover plates of the guide blade tips facing the channel 16 formed.
  • the blade roots are now provided with a concave contour and / or the guide blade tips are provided with a convex contour. Details of this contouring can be seen in FIGS. 6 and 7. 7 shows the concave contour 24b of a blade root 22.
  • the concave contour 24b of the blade root according to the invention is curved in the direction of the rotor and at this point increases the channel average.
  • the minimum of the concave contour 24b is offset in the direction of the rotor by the distance h 2H1 with respect to the base line G of the non-contoured channel.
  • the concave contour 24b is offset in the region of the front edge of the rotor blade against the base line G of the uncontoured channel by the distance h 2H2 in the direction of the rotor.
  • the value of k ' 2 is particularly preferably between 0.01 and 0.03.
  • the minimum of the concave contour 24b is offset downstream from the leading edge of the airfoil by the distance S 2H .
  • the value of k ' 6 is particularly preferably between 0.3 and 0.4.
  • the embodiment of the straight contour 28 of the blade tip 20 according to the invention shown in FIG. 7 has an angle of inclination ⁇ with respect to the base line G of the uncontoured blade tip, which is tapered here.
  • the relationship preferably applies to this angle of inclination 0 ° ⁇ ⁇ 4 °.
  • ⁇ is particularly preferably between 0 ° and 2 °.
  • FIG. 6 shows details of the concave contour 24a of a guide blade root 18.
  • the concave contour 24a of the rotor blade root according to the invention is curved in the direction of the stator and increases the channel average at this point.
  • the minimum of the concave contour 24a is offset in the direction of the stator by the distance h 1T1 in relation to that part of the base line G of the uncontoured channel located in the region of the front edge of the guide vane blade .
  • the value of k 1 is particularly preferably between 0.06 and 0.12.
  • the concave contour 24a is offset in the region of the front edge of the guide blade against the base line G of the uncontoured channel by the distance h 1T2 in the direction of the stator.
  • the value of k 2 is particularly preferably between 0.01 and 0.03.
  • the minimum of the concave contour 24a is offset downstream from the leading edge of the guide vane by the distance S 1T1 .
  • the value of k 3 is particularly preferably between 0.23 and 0.27.
  • the minimum of the concave contour of the guide vane root has an extensive flattened area of the width S 1T2 .
  • the value of k 4 is particularly preferably between 0.12 and 0.13.
  • FIG. 6 shows details of a convex contour 26 of a guide vane tip 16 according to the invention.
  • the convex contour 26 bulges into the flow-through channel and narrows the channel average.
  • the maximum of the convex contour is offset from the base line G by the distance h 1H .
  • the value of k 5 is particularly preferably between 0.04 and 0.08.
  • the maximum of the convex contour is offset downstream from the leading edge of the guide vane by the distance S 1H .
  • the value of k 6 is particularly preferably between 0.3 and 0.4.
  • the small arrows shown in Fig. 1 indicate the flow directions of the leak or Maze flows. Step on the rotor-side wall 8 of the flow channel they enter the labyrinth at 54, pass the tips of the shovels and flow in 56 back into the main flow. On the stator side flow limiting Wall 10 is the labyrinth entrance at 44 and the labyrinth exit at 46 designated. The return of the labyrinth flow to the main flow should be as possible take place without vertebrae. On the one hand, this is due to the configuration according to the invention the concave contour 24a of the guide blade root 18 is reached. As before This is mentioned above in the area of the front edge of the guide vane blade compared to the essentially conical baseline shown in FIG.
  • the upstream tip of the blade is essentially parallel to the mainstream.
  • the blade tip 20 the angle of inclination of the straight contour 28 preferably somewhat flatter than the angle of inclination of the conical base line of the stator-side flow-limiting Wall chosen. This results in a rectification of Maze and main flow.
  • the resetting of the concave contour 24a in The area of the front edge of the guide vane blade is particularly preferably selected such that that parts of the main flow also flow into the recess.
  • the return of the labyrinth flow on the rotor-side wall of the flow Channel is done analogously.
  • the concave contour 24b of the blade root is also here in the area of the front edge of the blade compared to that in the essential cylindrical baseline of the non-contoured channel towards the rotor set back. Both preferably flows into the recess formed thereby the labyrinth flow as well as part of the main flow.
  • the downstream one The area of the convex contour 26 is designed so that a Main and labyrinth current are rectified.
  • Entries 44 and 54 reduce the occurrence of labyrinth currents.
  • the angle of inclination of the downstream part of the concave contour 24b in Fig. 1 indicated by a dashed line, is preferably steeper than the angle of inclination of the upstream part of the convex contour 26. This causes a pressure reduction at the labyrinth inlet.
  • the one flowing through the labyrinth In a first approximation, current is proportional to that between the entrance of the labyrinth and Maze exit prevailing pressure difference. This results in the reduction the pressure at the labyrinth inlet in a reduction of the labyrinth current.
  • the pressure is also reduced at the labyrinth inlet 44.
  • the recesses in the concave contours 24a and 24b of the blade feet cause a shift in the position of highest load from the front to the rear area of the shovel. This causes a reduction in the flow rate on the suction side of the bucket and thus leads to a lowering of secondary losses.
  • FIGS. 2-5 show the velocity distribution of a mass particle along a Cut through the airfoil (guide blade, moving blade). The speed the mass particle was directly on the surface of the airfoil measured.
  • Each figure contains two pairs of curves, once two measurement curves along the suction or pressure side of an airfoil in a flow Channel with a straight contour, once corresponding measurement curves along a Blade in a flow-through channel with a particularly preferred contour according to the invention i.e. with concave contours 24a and 24b on the guide vane feet and on the blade roots, convex contour 26 on the guide blade tips and conical contour 28 at the blade tips.
  • the stator side of the flow through the channel conical and the rotor side of the flow channel is cylindrical.
  • FIG. 1 Each figure shows the speed distribution in a direct comparison along a certain height of the airfoil; in Figure 2 along of the guide vane root, in Fig. 3 along the blade tip, in Fig. 4 along the vane tip and in Fig. 5 along the blade root.
  • the little ones Measuring points shown in circles were in the channel with a straight contour and the measuring points represented by small squares in the invention contoured channel measured.
  • the upper curve branches (the two parts of the graphs in the respective figures the mostly higher Mach numbers than the other two parts of the graphs the measured values along the suction side (convex side) of the Blade while the lower-lying curves (lower curve branches) the measured values point along the pressure side (concave side) of the airfoil.
  • the horizontal coordinate axis (x-axis) of FIGS. 2-5 is the normalized one Arc length of an airfoil applied.
  • the x value denotes 0.0 the stagnation point of the airfoil.
  • the x value 1.0 indicates the furthest downstream outflow point of the airfoil.
  • the x value denotes 0.5 the point on the surface of the airfoil which is half of the total arc length corresponds, with the total arc length of the suction side usually is greater than the total sheet length of the printed page.
  • the Mach number results from the ratio of Velocity of the mass particle relative to the speed of sound in the working medium.
  • Figures 2-5 can show the degree of load absorption in the rear blade area be removed.
  • the normalized arc length can be the difference between the two Mach numbers specified for this value, i.e. the Mach number of the suction and pressure side.
  • the maximum of the Mach number difference provides information about the location of the maximum power transmission from Working medium on the shovel. It is worth striving for this maximum if possible far back, i.e. to shift to large x values in FIGS.
  • the recesses in the concave contours 24a and 24b of the blade feet cause a shift in the position of highest load from the front to the rear area of the shovel. This causes a reduction in the flow rate on the suction side of the bucket and thus leads to a lowering of secondary losses.
  • Fig. 2 shows that the maximum curve belly of about 0.43 was shifted about 0.5 backwards. This means an improvement the degree of load absorption in the rear area of the guide blade root.
  • Fig. 5 shows the blade root. Here the degree of load absorption became clear moved to the rear of the blade root.

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  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP98810233A 1998-03-19 1998-03-19 Veine profilée pour une turbomachine axiale Withdrawn EP0943784A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP98810233A EP0943784A1 (fr) 1998-03-19 1998-03-19 Veine profilée pour une turbomachine axiale

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Application Number Priority Date Filing Date Title
EP98810233A EP0943784A1 (fr) 1998-03-19 1998-03-19 Veine profilée pour une turbomachine axiale

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Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2001044623A1 (fr) * 1999-12-16 2001-06-21 Atlas Copco Tools Ab Machine a rotor de type turbine a flux axial destinee a fonctionner par fluide elastique
JP2002122002A (ja) * 2000-08-21 2002-04-26 General Electric Co <Ge> ロータ組立体の円周方向リム応力を減少させるための方法及び装置
EP1126132A3 (fr) * 2000-02-18 2003-05-02 General Electric Company Paroi radiale profilée pour compresseur
US6575693B2 (en) 2000-06-23 2003-06-10 Alstom (Switzerland) Ltd Labyrinth seal for rotating shaft
JP2008274926A (ja) * 2007-04-27 2008-11-13 Honda Motor Co Ltd 軸流型ガスタービンエンジンのガス通路形状
DE102008031789A1 (de) * 2008-07-04 2010-01-07 Deutsches Zentrum für Luft- und Raumfahrt e.V. Verfahren und Vorrichtung zur Beeinflussung von Sekundärströmungen bei einer Turbomaschine
DE102010027588A1 (de) * 2010-07-19 2012-01-19 Rolls-Royce Deutschland Ltd & Co Kg Fan-Nachleitradschaufel eines Turbofantriebwerks
DE102011076804A1 (de) * 2011-05-31 2012-12-06 Honda Motor Co., Ltd. Innenumfangsflächenform eines Axialverdichtergehäuses
EP3043022A1 (fr) 2014-12-12 2016-07-13 MTU Aero Engines GmbH Turbomachine a elargissement d'espace annulaire et aube
FR3106626A1 (fr) * 2020-01-24 2021-07-30 Safran Aircraft Engines Basculement differencié entre rotor et stator aux entrefers rotor-stator dans un compresseur de turbomachine
DE102021109844A1 (de) 2021-04-19 2022-10-20 MTU Aero Engines AG Gasturbinen-Schaufelanordnung

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2392673A (en) * 1943-08-27 1946-01-08 Gen Electric Elastic fluid turbine
US2846137A (en) * 1955-06-03 1958-08-05 Gen Electric Construction for axial-flow turbomachinery
US4606700A (en) * 1979-10-15 1986-08-19 Vsesojuzny Naucho-Issledovatelsky Institut Burovoi Tekhniki Turbodrill multistage turbine
US4832567A (en) * 1981-01-05 1989-05-23 Alsthom-Atlantique Turbine stage
US5626462A (en) * 1995-01-03 1997-05-06 General Electric Company Double-wall airfoil
EP0799973A1 (fr) 1996-04-01 1997-10-08 Asea Brown Boveri Ag Contour de paroi pour une turbomachine axiale
EP0846867A2 (fr) * 1996-12-06 1998-06-10 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Turbomachine avec un étage de compression transsonique

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2392673A (en) * 1943-08-27 1946-01-08 Gen Electric Elastic fluid turbine
US2846137A (en) * 1955-06-03 1958-08-05 Gen Electric Construction for axial-flow turbomachinery
US4606700A (en) * 1979-10-15 1986-08-19 Vsesojuzny Naucho-Issledovatelsky Institut Burovoi Tekhniki Turbodrill multistage turbine
US4832567A (en) * 1981-01-05 1989-05-23 Alsthom-Atlantique Turbine stage
US5626462A (en) * 1995-01-03 1997-05-06 General Electric Company Double-wall airfoil
EP0799973A1 (fr) 1996-04-01 1997-10-08 Asea Brown Boveri Ag Contour de paroi pour une turbomachine axiale
EP0846867A2 (fr) * 1996-12-06 1998-06-10 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Turbomachine avec un étage de compression transsonique

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
WALTER TRAUPEL: "Thermische Turbomaschinen", vol. 1, SPRINGER VERLAG 1966, article "Kapitel 7"

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6705834B1 (en) 1999-12-16 2004-03-16 Atlas Copco Tools Ab Axial flow turbine type rotor machine for elastic fluid operation
WO2001044623A1 (fr) * 1999-12-16 2001-06-21 Atlas Copco Tools Ab Machine a rotor de type turbine a flux axial destinee a fonctionner par fluide elastique
EP1126132A3 (fr) * 2000-02-18 2003-05-02 General Electric Company Paroi radiale profilée pour compresseur
US6575693B2 (en) 2000-06-23 2003-06-10 Alstom (Switzerland) Ltd Labyrinth seal for rotating shaft
JP4636746B2 (ja) * 2000-08-21 2011-02-23 ゼネラル・エレクトリック・カンパニイ ロータ組立体の円周方向リム応力を減少させるための方法及び装置
JP2002122002A (ja) * 2000-08-21 2002-04-26 General Electric Co <Ge> ロータ組立体の円周方向リム応力を減少させるための方法及び装置
EP1182328A3 (fr) * 2000-08-21 2003-06-04 General Electric Company Méthode pour réduire la tension circonférencielle dans des rotors
US8192154B2 (en) 2007-04-27 2012-06-05 Honda Motor Co., Ltd. Shape of gas passage in axial-flow gas turbine engine
JP2008274926A (ja) * 2007-04-27 2008-11-13 Honda Motor Co Ltd 軸流型ガスタービンエンジンのガス通路形状
JP2012127354A (ja) * 2007-04-27 2012-07-05 Honda Motor Co Ltd 軸流型ガスタービンエンジンのガス通路形状
DE102008031789A1 (de) * 2008-07-04 2010-01-07 Deutsches Zentrum für Luft- und Raumfahrt e.V. Verfahren und Vorrichtung zur Beeinflussung von Sekundärströmungen bei einer Turbomaschine
DE102010027588A1 (de) * 2010-07-19 2012-01-19 Rolls-Royce Deutschland Ltd & Co Kg Fan-Nachleitradschaufel eines Turbofantriebwerks
US8784042B2 (en) 2010-07-19 2014-07-22 Rolls-Royce Deutschland Ltd & Co Kg Fan downstream guide vanes of a turbofan engine
DE102011076804B4 (de) 2011-05-31 2019-04-25 Honda Motor Co., Ltd. Innenumfangsflächenform eines Lüftergehäuses eines Axialverdichters
DE102011076804A1 (de) * 2011-05-31 2012-12-06 Honda Motor Co., Ltd. Innenumfangsflächenform eines Axialverdichtergehäuses
EP3043022A1 (fr) 2014-12-12 2016-07-13 MTU Aero Engines GmbH Turbomachine a elargissement d'espace annulaire et aube
DE102014225689A1 (de) 2014-12-12 2016-07-14 MTU Aero Engines AG Strömungsmaschine mit Ringraumerweiterung und Schaufel
US10570743B2 (en) 2014-12-12 2020-02-25 MTU Aero Engines AG Turbomachine having an annulus enlargment and airfoil
FR3106626A1 (fr) * 2020-01-24 2021-07-30 Safran Aircraft Engines Basculement differencié entre rotor et stator aux entrefers rotor-stator dans un compresseur de turbomachine
DE102021109844A1 (de) 2021-04-19 2022-10-20 MTU Aero Engines AG Gasturbinen-Schaufelanordnung
US11585223B2 (en) 2021-04-19 2023-02-21 MTU Aero Engines AG Gas turbine blade arrangement

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