EP2204542A2 - Configuration de pied d'aube de turbine inclinée - Google Patents

Configuration de pied d'aube de turbine inclinée Download PDF

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Publication number
EP2204542A2
EP2204542A2 EP09178822A EP09178822A EP2204542A2 EP 2204542 A2 EP2204542 A2 EP 2204542A2 EP 09178822 A EP09178822 A EP 09178822A EP 09178822 A EP09178822 A EP 09178822A EP 2204542 A2 EP2204542 A2 EP 2204542A2
Authority
EP
European Patent Office
Prior art keywords
rotor blade
root
dovetail
suction
turbine
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Withdrawn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
EP09178822A
Other languages
German (de)
English (en)
Other versions
EP2204542A3 (fr
Inventor
Bradley T. Boyer
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 Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP2204542A2 publication Critical patent/EP2204542A2/fr
Publication of EP2204542A3 publication Critical patent/EP2204542A3/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Withdrawn legal-status Critical Current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
    • F01D5/34—Rotor-blade aggregates of unitary construction, e.g. formed of sheet laminae
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • 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
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • 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
    • F05D2240/80—Platforms for stationary or moving blades

Definitions

  • This present application relates generally to apparatus, methods and/or systems concerning improved turbine blade root configurations. More particularly, but not by way of limitation, the present application relates to apparatus, methods and/or systems pertaining to turbine blades that combine axial entry, linear dovetails with curved platforms.
  • nested is a common term that refers to a condition wherein the curvature of neighboring airfoils overlaps. This overlap generally means that the turbine blades, if aligned as they might be when installed in a rotor wheel of a conventional turbine engine, cannot be separated with an axial or a linear movement of one of the blades because of the interference between the nested airfoils, i.e., the airfoils would make contact and prevent separation in this manner.
  • the present application thus describes a rotor blade for use in a turbine engine, the rotor blade comprising a root and, extending in a radial direction from the root, an airfoil, wherein the root includes at least one root aligned surface that is tilted. Tilted may comprises a non-radial orientation.
  • the aligned surfaces may comprise the surfaces along the root that are configured to align with and be relatively closely spaced from or in contact with the aligned surfaces of the root of a neighboring rotor blade.
  • the rotor blade may comprise at least two root aligned surfaces, one of which resides on a pressure side of the rotor blade and the other of which resides on a suction side of the rotor blade. All of the root aligned surfaces may be tilted.
  • Figure 1 illustrates a schematic representation of a gas turbine engine 100.
  • gas turbine engines operate by extracting energy from a pressurized flow of hot gas that is produced by the combustion of a fuel in a stream of compressed air.
  • gas turbine engine 100 may be configured with an axial compressor 106 that is mechanically coupled by a common shaft to a downstream turbine section or turbine 110, and a combustor 112 positioned between the compressor 106 and the turbine 110.
  • the following invention may be used in all types of turbine engines, including, for example, gas turbine engines, steam turbine engines, and aircraft engines.
  • the invention will be described in relation to a gas turbine engine, though this description is exemplary only and not intended to be limiting in any way.
  • Figure 2 illustrates a view of an exemplary multi-staged axial compressor 118 that may be used in a gas turbine engine.
  • the compressor 118 may include a plurality of stages. Each stage may include a row of compressor rotor blades 120 followed by a row of compressor stator blades 122.
  • a first stage may include a row of compressor rotor blades 120, which rotate about a central shaft, followed by a row of compressor stator blades 122, which remain stationary during operation.
  • the compressor stator blades 122 generally are circumferentially spaced one from the other and fixed about the axis of rotation.
  • the compressor rotor blades 120 are circumferentially spaced and attached to the shaft such that, when the shaft rotates during operation, the compressor rotor blades 120 rotates about it.
  • the compressor rotor blades 120 are configured such that, when spun about the shaft, they impart kinetic energy to the air or working fluid flowing through the compressor 118.
  • the compressor 118 may have many other stages beyond the stages that are illustrated in Figure 2 . Additional stages may include a plurality of circumferential spaced compressor rotor blades 120 followed by a plurality of circumferentially spaced compressor stator blades 122.
  • FIG 3 illustrates a partial view of an exemplary turbine section or turbine 124 that may be used in the gas turbine engine.
  • the turbine 124 also may include a plurality of stages. Three exemplary stages are illustrated, but more or less stages may present in the turbine 124.
  • Each stage may include a plurality of turbine buckets or turbine rotor blades 126, which rotate about the shaft during operation, and a plurality of nozzles or turbine stator blades 128, which remain stationary during operation.
  • the turbine stator blades 128 generally are circumferentially spaced one from the other and fixed about the axis of rotation.
  • the turbine rotor blades 126 may be mounted on a turbine wheel (not shown) for rotation about the shaft (not shown).
  • each additional stage may include a row of turbine stator blades 128 followed by a row of turbine rotor blades 126.
  • rotor blades is a reference to the rotating blades of either the compressor 118 or the turbine 124, which include both compressor rotor blades 120 and turbine rotor blades 126.
  • stator blades is a reference to the stationary blades of either the compressor 118 or the turbine 124, which include both compressor stator blades 122 and turbine stator blades 128.
  • blades will be used herein to refer to either type of blade.
  • blades is inclusive to all type of turbine engine blades, including compressor rotor blades 120, compressor stator blades 122, turbine rotor blades 126, and turbine stator blades 128.
  • the rotation of compressor rotor blades 120 within the axial compressor 118 may compress a flow of air.
  • energy may be released when the compressed air is mixed with a fuel and ignited.
  • the resulting flow of hot gases from the combustor 112 then may be directed over the turbine rotor blades 126, which may induce the rotation of the turbine rotor blades 126 about the shaft, thus transforming the energy of the hot flow of gases into the mechanical energy of the rotating blades and, because of the connection between the rotor blades in the shaft, the rotating shaft.
  • the mechanical energy of the shaft may then be used to drive the rotation of the compressor rotor blades 120, such that the necessary supply of compressed air is produced, and also, for example, a generator to produce electricity.
  • FIG 4 depicts a portion of a turbine assembly 130 of the gas turbine engine 100.
  • the turbine assembly 130 may be mounted downstream from the combustor (not shown in Figure 4 ) for receiving hot combustion gases 131 therefrom.
  • the turbine assembly 130 generally comprises a disk 132 having a plurality of turbine rotor blades 126 securely attached thereto.
  • the turbine rotor blade 126 comprises an airfoil 136 that extends radially from a root 138, which it generally is integral therewith.
  • a platform 140 is disposed at the base of the airfoil 136 and generally is also integral therewith.
  • the turbine assembly 130 is axisymmetrical about an axial centerline axis 141.
  • An annular shroud 142 surrounds the blades 126 and is suitably joined to a stationary stator casing (not shown).
  • the shroud 142 provides a relatively small clearance or gap between it and the rotor blades 126, which limits the leakage of combustion gases 131 over the blades 126 during operation.
  • the airfoil 136 generally includes a concave pressure sidewall or pressure side 143 and a circumferentially or laterally opposite, convex suction sidewall or suction side 144. Both the pressure sidewall 143 and the suction sidewall 144 extend axially between a leading edge 146 and a trailing edge 148. The pressure sidewall 143 and the suction sidewall 144 further extend in the radial direction between the radially inner root 138 at the platform 140 and a radially outer blade tip 150.
  • the root 138 generally includes a shank 152, the outer radial surface of which is the platform 140, and a dovetail 154.
  • the dovetail 154 is the inner radial section of the root 138, while the shank 152 is the section that connects the dovetail 154 to the airfoil 136.
  • the dovetail 154 has a side entry type configuration that includes a plurality of tangs 156, which generally provides the root 138 with a serrated cross-section.
  • the shank 152 extends from the outer radial portion of the dovetail 154 to the outer radial surface of the shank 152, which, as stated, is the platform 140.
  • the root 138 may be described as having a trailing edge or face 158 and a leading edge or face 160, and, as illustrated, the root 138 may extend in a linear direction from the trailing face 158 to the leading face 160.
  • the root 138 may be described as having a pressure face 162 and a suction face 164, which correspond, respectively, with the pressure side 143 and the suction side 144 of the airfoil 136.
  • the disc 132 may have a plurality of dovetail grooves 166 formed around its circumference.
  • Each of the dovetail grooves 166 may be formed as a mate to the dovetails 154 of the rotor blades 126 such that each of the dovetails 154 may be axially inserted into the dovetail groove 162. It will be appreciated that the configuration of the dovetail 154/dovetail groove 166 connects the rotor blades 126 to the disc 132 and prevents the radial displacement of the rotor blades 126 during operation.
  • the dovetail 154 may be linear, i.e., have a linear orientation from the trailing face 158 to the leading face 160, and the dovetail groove 162 may be linearly oriented as well. Formed in this manner, the rotor blades 126 may be axially inserted into the dovetail grooves 162 a linear fashion. As discussed in more detail below, a curved configuration for the root is also possible.
  • Turbine rotor blades are the rotating blades within the turbine section of the turbine engine. This description is exemplary only, as embodiments of the invention described herein are not limited to usage with only turbine rotor blades.
  • the present invention also may be applied to compressor rotor blades, which, generally, are the rotating blades within the compressor section of the turbine engine. Accordingly, reference herein to "rotor blades,” without further specificity, is meant to be inclusive of both turbine rotor blades and compressor rotor blades. And, for instance, examples that are applied to turbine rotor blades are not meant to exclude usage of the present invention in compressor rotor blades.
  • Figure 5 depicts a rotor blade with a conventional linear root 138.
  • the linear root 138 includes a platform 140 and a dovetail 154 that have a linear orientation from the trailing face 158 to the leading face 160 of the root 138. More particularly, the pressure face 162 and the suction face 164 of the root 138 are not curved and generally run in a straight from the trailing face 158 to the leading face 160. It will be appreciated that the linearly oriented platform 140 is approximately rectilinear in shape. Each edge of the platform 140 may be identified by its relationship to the trailing face 158, leading face 160, the pressure face 162, and the suction face 164.
  • the platform 140 may be described to include a trailing edge 170, a leading edge 172, a pressure edge 174, and a suction edge 176.
  • the pressure edge 174 is generally linear or straight.
  • the suction edge 176 is generally linear or straight.
  • the dovetail 154 also may extend from the trailing face 158 to the leading face 160 in an approximately linear manner. Other portions of the shank 152 also may be linear.
  • performance criteria for airfoil design may require that airfoils become "nested" when positioned in an assembled configuration. When this is the case, removing blades linearly (which is what would be the case with linear configurations similar to Figure 5 ) becomes impossible.
  • Figure 6 depicts a rotor blade with a conventional curved root 138.
  • the curved root may include a curved platform 140 and a curved dovetail 154.
  • the pressure face 162 and the suction face 164 of the root 138 are curved.
  • the pressure edge 174 of the platform 140 may form a concave curve.
  • the suction edge 176 of the platform 140 may form a similar curve, though it may be a convex curve.
  • the dovetail 154 also may form a similar curve.
  • Other portions of the shank 152 may form a similar curve.
  • the curvature for all of these components may be similar and, generally, is an arc of a circle.
  • Figure 7 depicts a rotor blade with a curved platform 140 and a linear dovetail 154.
  • the dovetail 154 may be substantially similar to the dovetail 154 of Figure 5 . That is, the dovetail 154 may be substantially linear and be configured to mate with a substantially linear dovetail groove 166.
  • the linear dovetail 154 and dovetail groove 166 may be aligned such that, on installation, each runs parallel with the centerline axis 141. In other cases, the linear dovetail 154 and the dovetail groove 166 may be skewed in relation to the direction of the centerline axis 141.
  • the platform 140 may be curved, i.e., substantially similar to the platform 140 configuration of Figure 6 .
  • the pressure edge 174 of the platform 140 may form a curve, which may be a concave curve.
  • the suction edge 176 of the platform 140 may form a similar curve, though the suction edge 176 may form a convex curve.
  • the curvature of the suction edge 176 and the pressure edge 174 may be substantially the same, though offset by the width of the platform 140. In this manner, the pressure edge 174 of one blade may engage the suction edge 176 of a neighboring blade so that the platform 140 of the neighboring blades forms a smooth substantially continuous surface.
  • the trailing edge 170 and the leading edge 172 of the platform 140 may remain linear, though this is not required.
  • the portions of the shank 152 below the platform generally may form a transition between the curved platform 140 and the linear dovetail 154.
  • the curvature of the pressure edge 174 and the suction edge 176 may be approximately the same.
  • the curve of the pressure edge 174 and the suction edge 176 may form the arc of an approximate circle.
  • root configurations consistent with the present invention may provide advantages associated linear root configurations, such as the one illustrated in Figure 5 , while also providing advantages associated with curved root configurations, such as the one illustrated in Figure 6 .
  • adjacent rotor blades as they are typically configured in an installed position on a rotor wheel, have aligned surfaces that are adjacent or separated by a relatively small distance. These surfaces, which, for the sake of brevity, will be referred to herein as root aligned surfaces or "aligned surfaces 178", are so closely spaced apart that they appear to rest against one another. As one of ordinary skill in the art will appreciate, however, generally these surfaces are separated by a very narrow space and do not make contact with one another, though, in certain applications, contact between the two surfaces is possible.
  • root aligned surfaces or “aligned surfaces” refers to any of the surfaces along the root of a rotor blade that are aligned with and very closely spaced from or, in some instances, in contact with the aligned surfaces of the root of a neighboring rotor blade.
  • the aligned surfaces 178 generally are configured such that the apparent junction between the two surfaces is made across opposing approximately planar lateral surfaces, with the narrow gap defined therebetween, when viewed in cross-section, forming an apparent junction line 181 that is substantially linear.
  • the junction line 181 formed between the generally planar surfaces of opposing root aligned surfaces 178 is a substantially radially oriented line, i.e., a line that forms an approximate 0° angle with a line extending from the axis of the turbine in a radial direction (i.e., perpendicular to the axis of the turbine).
  • the adjacent rotor blades 126 have several aligned surfaces 178 along their root portions 138.
  • the root aligned surfaces 178 may include the pressure edge 174 of a first rotor blade 126 aligning with and being adjacent to the opposing suction edge 176 of a second (and neighboring) rotor blade 126, as well as the suction edge 176 of the first rotor blade aligning with and being adjacent to the pressure edge 174 of a third (and also neighboring rotor blade 126).
  • root aligned surfaces 178 may include the opposing sides of coverplates 180 that may be formed on adjacent rotor blades 126, as further shown in Figure 8 .
  • the trailing face 158 and/or leading face 160 of the shank 152 may be substantially "covered” or enclosed by a coverplate 180, which generally comprises a relatively thin rectangular plate.
  • a coverplate 180 which generally comprises a relatively thin rectangular plate.
  • the junction line 181 that is formed between any of these exemplary root aligned surfaces 178 is substantially radially oriented and perpendicular with the axis of the turbine, i.e., if the junction line 181 were extended, it would substantially intersect the axis of the turbine and be approximately perpendicular therewith.
  • the opposing root aligned surfaces 178 may be configured such that the junction line 181 is tilted, i.e., not radially oriented. As prescribed herein, the junction line 181 may form an angle with a radially oriented line 183, with this tilting providing certain operational advantages.
  • Figure 9 illustrates several installed turbine rotor blades 126, each including an airfoil 136 and a root 138, wherein the configuration of the root 138 is consistent with the current invention. Similar to other descriptions herein, the root 138 of Figure 9 includes a shank 152 with an outer radial platform 140 and a dovetail 154.
  • the trailing face 158 and/or leading face 160 of the shank 152 may be substantially “covered” or enclosed by a coverplate 180 (or, as depicted in Figure 10 , the trailing face 158 and/or leading face 160 of the shank 152 may be "uncovered”).
  • coverplates are driven by several operational criteria and that the invention described herein is applicable whether or not the coverplates are included in the design of the shank. Further, coverplates may be integral to the shank or attached thereto, neither of which affect the usage or applicability of the present invention.
  • the root aligned surfaces 178 of the adjacent rotor blades 126 include the aligned and adjacent surfaces between the pressure edge 174 and the suction edge 176 and the sides of the coverplates 180, as illustrated. Consistent with the present invention, the root aligned surfaces 178 are configured such that the junction line 181 between them forms an angle ⁇ with the radially oriented line 183. That is, angle ⁇ represents the approximate angle between 1) the approximate junction line 181 that is formed between the opposing root aligned surfaces 178 and 2) a radially orient line 183 (i.e., a line that approximately intersects and is perpendicular to the axis of the turbine engine). In some embodiments, the angle ⁇ is between approximately 0° and 60°. More preferably, the angle ⁇ is between approximately 15° and 45°. More preferably still, the angle ⁇ is between approximately 25° and 35°. And, ideally, the angle ⁇ is approximately 30°.
  • the root aligned surfaces 178 no longer include the side surfaces of the coverplates 180 and, thus, primarily consist of the aligned and adjacent surfaces between the opposing pressure faces 162 and the suction faces 164 of adjacent rotor blades 126.
  • the pressure face 162 of a first rotor blade 126 may align with the suction face 164 of a second rotor blade 126 that is adjacent to the first rotor blade 126
  • the suction face 164 of the first rotor blade 126 may align with the pressure face 162 of a third rotor blade 126 that also is adjacent to the first rotor blade 126.
  • the root aligned surfaces 178 along the pressure face 162 and the suction face 164 generally are limited to the surface areas along the pressure edge 174 and the suction edge 176. Whatever the case may be (i.e., whatever form the aligned surfaces 178 between the roots 138 of adjacent rotor blades take), the root aligned surfaces 178 may be configured such that the junction line 181 formed therebetween forms an angle ⁇ with a radially oriented line 183 in the manner described above.
  • tilted root aligned surfaces in accordance with the invention described herein.
  • this type of geometry is beneficial to certain turbine blade attachment geometries, particularly those involving high chord, high camber airfoils that have short shanks and skewed axial entry dovetails.
  • One advantage of this design is that it allows the blade geometry to include an integral coverplate that creates a continuous surface of revolution on the forward and/or aft vertical faces of the shank area.
  • the non-radial angle (i.e., angle 0) also creates greater, more uniform aligned surface between seal pins and rotor blades, which, among other advantages, reduces leakage and thereby improves efficiency.
  • the current invention is applicable to turbine blades that have a curved platform and a straight or linear dovetail configuration, such as those described above in relation to Figures 1-7 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP09178822A 2008-12-30 2009-12-11 Configuration de pied d'aube de turbine inclinée Withdrawn EP2204542A3 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/346,301 US20100166561A1 (en) 2008-12-30 2008-12-30 Turbine blade root configurations

Publications (2)

Publication Number Publication Date
EP2204542A2 true EP2204542A2 (fr) 2010-07-07
EP2204542A3 EP2204542A3 (fr) 2013-04-03

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP09178822A Withdrawn EP2204542A3 (fr) 2008-12-30 2009-12-11 Configuration de pied d'aube de turbine inclinée

Country Status (4)

Country Link
US (1) US20100166561A1 (fr)
EP (1) EP2204542A3 (fr)
KR (1) KR20100080451A (fr)
CN (1) CN101793168A (fr)

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Publication number Priority date Publication date Assignee Title
FR3045709A1 (fr) * 2015-12-21 2017-06-23 Snecma Aube de soufflante
WO2017209752A1 (fr) * 2016-06-02 2017-12-07 Siemens Aktiengesellschaft Système de fixation asymétrique pour aube de turbine
FR3107551A1 (fr) * 2020-02-20 2021-08-27 Safran Aircraft Engines Aube de turbine

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US8007245B2 (en) * 2007-11-29 2011-08-30 General Electric Company Shank shape for a turbine blade and turbine incorporating the same
FR2944050B1 (fr) * 2009-04-02 2014-07-11 Turbomeca Roue de turbine a pales desaccordees comportant un dispositif d'amortissement
US20120156045A1 (en) * 2010-12-17 2012-06-21 General Electric Company Methods, systems and apparatus relating to root and platform configurations for turbine rotor blades
US10036261B2 (en) * 2012-04-30 2018-07-31 United Technologies Corporation Blade dovetail bottom
US9470098B2 (en) * 2013-03-15 2016-10-18 General Electric Company Axial compressor and method for controlling stage-to-stage leakage therein
US10830065B2 (en) * 2015-06-02 2020-11-10 Siemens Aktiengesellschaft Attachment system for a turbine airfoil usable in a gas turbine engine
WO2019046006A1 (fr) * 2017-08-28 2019-03-07 Siemens Aktiengesellschaft Plates-formes à géométrie avancée pour aubes de turbine
KR20220165474A (ko) 2021-06-08 2022-12-15 이희두 버추얼리티(vr) 기반의 모바일 커머스 토탈 서비스 플랫폼
KR20220166111A (ko) 2021-06-09 2022-12-16 이희두 버추얼리티(vr) 기반의 전통시장 모바일 커머스 서비스 플랫폼
CN114263632A (zh) * 2021-10-22 2022-04-01 中国航发沈阳发动机研究所 一种发动机风扇转子部件

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US3545882A (en) * 1968-01-17 1970-12-08 Rolls Royce Pressure exchanger rotor
DE1953709A1 (de) * 1968-10-28 1970-04-30 Elin Union Ag Beschaufelung der Rotoren von Dampf- oder Gasturbinen
JPS5493702A (en) * 1977-12-28 1979-07-25 Kawasaki Heavy Ind Ltd Rotor for multistage axial-flow rotary machine
GB2139295A (en) * 1983-05-05 1984-11-07 Tuomo Kaivola Thermal joint e.g. for a turbine
JPH04123302U (ja) * 1991-04-22 1992-11-09 三菱重工業株式会社 回転翼
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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3045709A1 (fr) * 2015-12-21 2017-06-23 Snecma Aube de soufflante
WO2017209752A1 (fr) * 2016-06-02 2017-12-07 Siemens Aktiengesellschaft Système de fixation asymétrique pour aube de turbine
FR3107551A1 (fr) * 2020-02-20 2021-08-27 Safran Aircraft Engines Aube de turbine

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EP2204542A3 (fr) 2013-04-03
US20100166561A1 (en) 2010-07-01
KR20100080451A (ko) 2010-07-08

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