WO2019012102A1 - Pale de compresseur et compresseur axial comprenant ladite pale - Google Patents

Pale de compresseur et compresseur axial comprenant ladite pale Download PDF

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Publication number
WO2019012102A1
WO2019012102A1 PCT/EP2018/069072 EP2018069072W WO2019012102A1 WO 2019012102 A1 WO2019012102 A1 WO 2019012102A1 EP 2018069072 W EP2018069072 W EP 2018069072W WO 2019012102 A1 WO2019012102 A1 WO 2019012102A1
Authority
WO
WIPO (PCT)
Prior art keywords
airfoil
compressor
sided
blade
profile
Prior art date
Application number
PCT/EP2018/069072
Other languages
English (en)
Inventor
Christian Cornelius
Stephan Klumpp
David Monk
Original Assignee
Siemens Aktiengesellschaft
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 Siemens Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Publication of WO2019012102A1 publication Critical patent/WO2019012102A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • F04D29/544—Blade shapes
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/125—Fluid guiding means, e.g. vanes related to the tip of a stator vane
    • 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
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
    • 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

Definitions

  • Compressor blade and axial compressor comprising such
  • the invention relates to a compressor blade for an axial compressor, comprising a blade root for attaching the compressor blade to a carrier and a cantilevered airfoil extending from said blade root in a span direction from said blade root at 0% height to a free ending airfoil tip at 100% height, the airfoil comprising a suction side and a pressure side extend ⁇ ing in chord direction from an upstream-sided leading edge to a downstream-sided trailing edge, wherein the airfoil com ⁇ prises a tip-sided airfoil area extending from approximately 70 % height of the airfoil to the airfoil tip, wherein for each profile of the airfoil a stagger angle is determinable. Further the invention relates to an axial compressor with at least one ring of said compressor blades.
  • the object of the invention is the provision of a compressor blade and an axial compressor with increased efficiency.
  • the invention directs to cantilevered compressor blades suf- fering under tip losses.
  • the invention proposes to distribute the work con ⁇ version or the turning of the airfoil along the span direction in that way, that tip losses are reduced.
  • Compressor ro ⁇ tor blades shall perform more work conversion in the tip- sided airfoil area than in the remaining airfoil areas. For compressor stator vanes the turning is increased towards the tips.
  • the stagger angles of the airfoil are selected along the airfoil span such that compared to a conventional blading the flow turning is shifted partially from the tip- sided airfoil area into an intermediate airfoil area and from intermediate airfoil area to a root-sided airfoil area.
  • the work conversion rate in the tip-sided air ⁇ foil is decreased. This leads to the result, that the mass flow of air to be compressed is pushed away from that radial free end of the airfoil towards the endwall being a part of the blade root. Due to less main mass flow of air in the tip- sided airfoil area, resulting tip losses and endwall block ⁇ ages are reduced.
  • the invention proposes for a com ⁇ pressor blade as mentioned into the introduction, that the profile comprising the largest stagger angle of all profiles of the tip-sided airfoil area is located between 78% height and 86% height and that starting from the profile with the largest stagger angle of the tip-sided airfoil area the stag ⁇ ger angles of the remaining profiles of the tip-sided airfoil area decrease monotonously towards both 70 % height and the blade tip.
  • the invention is based on the knowledge that with the inven ⁇ tion the spanwise distribution of work conversion or turning is distributed in that way that more work conversion or turn ⁇ ing is established in the free ending tip region and less at the root-span of the airfoil. This reduces significantly losses within the row of compressor blades. Further endwall losses can be reduced by shifting the secondary flows. These features are very beneficial for compressor rotor blades as well as for compressor guide vanes. Hence the com ⁇ pressor blade could be embodied as a rotor blade or a stator vane while the aerodynamical effects are different, e.g. flow can be redistributed to one endwall only and/or flutter in ⁇ clination .
  • the airfoil comprises further a root-sided airfoil area ex ⁇ tending between the blade root and a first virtual separative profile located between 20% and 40% height, especially at 30% height, and an intermediate airfoil area extending between the first virtual separative profile and a second virtual separative profile located at said approximately 70% height.
  • the stagger angles in the root-sided airfoil area are approx ⁇ imately constant.
  • a stagger angle can be un ⁇ derstood as approximately constant, if its value does not change more than +/-2°. This increases the stiffness of the airfoil and enables a better monolithic attachment of the airfoil to the blade root for increasing lifetime of the com ⁇ pressor blade.
  • the profile with the smallest stagger angle of all profiles of the root-sided airfoil area is located between 17% height and 25% height and starting from said profile with the smallest stagger angle of the root-sided airfoil area
  • the stagger angles of the remaining profiles of the root- sided airfoil area increase monotonously towards both the first virtual separative profile and the blade root. Instead of increased stability and lifetime provides this an in ⁇ creased aerodynamical efficiency.
  • the stagger angles of the root-sided airfoil area increase continuously from the first virtual separative profile towards the root-sided profile for at least 1,5°, but not more than 7°. Such a decrease of the stagger angle from the blade root towards the first virtual separative profile enables a snooze transition of the pres ⁇ sure side surface and a suction side surface for redistrib ⁇ uting the local turning of the airfoil, which has been shown in detailed analysis and in comprehensive developments.
  • the stagger angles continuously increase towards to the profile section of the airfoil tip.
  • the airfoil comprises in span direction an over ⁇ all height and wherein the root-sided airfoil area and/or the tip-sided airfoil area comprises an extend parallel to the span, which is not larger than 30% of the overall height.
  • the stagger angle of the profiles of the intermedi ⁇ ate airfoil area changes monotonously from first virtual separative profile to the second virtual separative profile in a range between 4° and 8°.
  • Figure 1 shows a longitudinal section through the axial compressor of a gas turbine
  • Figure 2 shows a compressor rotor blade according to the invention as a first exemplary embodiment
  • Figure 3 shows schematically a graph for the stagger an ⁇ gle of a preferred compressor rotor blade
  • Figure 4 shows a compressor guide vane according to the invention as a second exemplary embodiment
  • Figure 5 shows schematically a graph for the stagger an ⁇ gle of a preferred compressor guide vane.
  • Figure 1 shows in a longitudinal section an axial compressor
  • Each compressor stage 11 comprises an upstream positioned compressor rotor blade 12 and a downstream positioned compressor stator vane 14.
  • the term compressor blade 42 is used.
  • the compressor rotor blades 12 are attached to the ro ⁇ tor 16, in example to a compressor rotor disk, in a conventional way.
  • the compressor stator vanes 14 are attached accordingly to compressor casing 18. Due to the schematic structure of the drawings the details of the attachment and especially the compressor blade roots are not shown.
  • Both the compressor rotor blades 12 and the compressor stator vanes 14 are embodied with cantilevered air ⁇ foils 20, 22 integrally attached to their roots, so that the airfoil 20 of the compressor rotor blade comprises a free ending tip 24, which faces the stationary compressor casing 18 under establishing a tip gap 26.
  • the airfoils 22 of the compressor stator vanes 14 comprise a tip 28 which faces the rotating rotor hub endwall 30 establishing a tip gap 32.
  • the rotor 16 rotates about its machine axis 17.
  • Figure 2 shows in a perspective view both a conventionally rotor blade 13 and a rotor blade 12 according to a first ex ⁇ emplary embodiment of the invention.
  • the conventional com- pressor rotor blade 13 is drawn in dashed line, while the compressor rotor blade 12 according to a first exemplary embodiment of the invention is drawn in an uninterrupted line.
  • Each compressor blade 42 comprises a blade root 19, from which the airfoil 20, 22 extend in span direction to the re ⁇ spective blade tip 24, 28.
  • the span axis is aligned with the radial direction R (figure 1) according to the machine axis 17.
  • the airfoil 20, 22 comprises a leading edge 25 and a trailing edge 27 between which a suction side 29 and a pres- sure side 31 extends in chord direction.
  • the airfoil height is determined at the leading edge 25 from its inner end at the blade root 19 with 0% height to its outer end at the blade tip 24, 28 with 100% height. For each height of the airfoils 20, 22 a profile can be de ⁇ termined.
  • the profile represents for a specific height of the airfoil the two-dimensional outer airfoil shape defined by a cross section through said airfoil at said height, the cross section being substantially parallel to the machine axis.
  • a stagger angle can be determined in a conven ⁇ tional way between a chord line of the respective profile and the compressor axial direction.
  • the chord line is the line between the points where the front and the rear of the pro- file would touch the surface, when a profile were laid convex side up on a flat surface.
  • stagger angles can be understood as those that turns a profile into a more closed position and smaller stag- ger angles can be understood as those that turns a profile into a more open position.
  • Stagger angles are usually in the range between 90° and the 180°.
  • the airfoil 20 is virtually separated in a number of airfoil areas, which follows successively one after one.
  • the compressor rotor blade 12 comprises three airfoil sections: a root-sided airfoil area 20a, an intermediate airfoil area 20b and a tip-sided airfoil area 20c.
  • the neighboring airfoil areas 20a, 20b, 20c are separated accordingly by a first virtual separative profile
  • the first virtual separative profile 21 is located at 30% height and the second virtual separative profile 23 is located at 70% height.
  • the tip-sided airfoil area 20c of the compressor rotor blade 12 comprises a profile 50 with the largest stagger angle of the tip-sided airfoil area 20c at about 82% height. With increasing distance to said height of 82% the stagger angles of the remaining profiles of the tip-sided airfoil area 20a decreases monotonously towards both the blade tip 24 and the second virtual separative profile 23.
  • the stagger an- gles of its profiles change monotonously very slightly from the first virtual separative profile 21 to the second virtual separative profile 23.
  • the change of the stagger angle within their interme ⁇ diate airfoil areas 20b, 22b is not larger than 3° from one end of said area to other end of said area and for compressor rotor blades the change is not larger than 12°.
  • the stagger angles of the root-sided airfoil area 20a increase continuously about 4° from the first virtual separative profile 21 towards the innermost root-sided pro ⁇ file at 0% height.
  • the profiles of the root-sided airfoil area 20a have stagger angles that are turned close in comparison to the stagger angle of the intermediate airfoil area 20b. This means that the leading edge 25 of the profiles are turned towards the suction side 29 while the trailing edges 27 are turned to the pressure side 31, also compared with the profiles of the intermediate airfoil area 20b, especially for compressor stator blades.
  • Figure 3 shows only schematically a graph 44 for the stagger angles for all profiles of the airfoil 20 of the compressor rotor blade 12 over to the span direction from 0% height to 100% height at the free ending airfoil tip, related to said first exemplary embodiment in full line.
  • the stagger angles decrease monoto ⁇ nously towards 0% height such that the smallest stagger angle is located approximately at 0% height.
  • the stagger angles remain approximately constant.
  • the stagger angles increase monotonously from the first virtual separation profile 21 towards the second virtual separation profile 23 with an approximately constant gradient.
  • the profile comprising the largest stagger angle of all profiles of the tip-sided airfoil area 20c is located between 78% height and 86% height and that starting from the profile with the largest stagger angle of the tip-sided airfoil area 20c the stagger angles of the remaining profiles of the tip-sided airfoil area 20c, 22c towards both the second virtual 70% height separative profile (23) and the blade tip decrease monotonously.
  • a dashed line 46 in figure 3 represents a variant of the graph for the stagger angles to the first preferred embodi ⁇ ment. Said variant differs from the first embodiment only in the root-sided airfoil area 20a. According to said variant in the root-sided airfoil area 20a the profile with the smallest stagger angle is located at 18% height. With increasing dis ⁇ tance to said height of 18% the stagger angles of the remain ⁇ ing profiles of the root-sided airfoil area 20a increases mo ⁇ notonously towards both the profile next to the blade root 19 at 0% height and the first virtual separative profile 21.
  • FIG 2 shows figure 4 both a convention ⁇ al compressor stator vane 15 and a compressor stator vane 14 according to the invention.
  • the conventional compressor stator vane 15 is drawn in dashed line, while the compressor stator vane 14 according to a second exemplary embodiment of the invention is drawn in an uninterrupted line.
  • Each compressor stator vane 14 comprises a blade root 19, from which the airfoil 22 extend in span direction to the respective blade tip 28.
  • the span axis is aligned with the ra ⁇ dial direction R (figure 1) according to the machine axis 17.
  • the airfoil 22 comprises a leading edge 25 and a trailing edge 27 between which a suction side 29 and a pressure side 31 extends in chord direction.
  • the airfoil height is deter ⁇ mined again at the leading edge 25 again from blade root 19 at 0% height to the blade tip 24 at 100% height.
  • each compressor rotor blade 12 of the axial compressor 10 is located on the inner diame- ter of the annular flow path 13 with respect to the machine axis 17, whereas the blade root 19 of each compressor stator vane 14 is located on the outer diameter of the annular flow path 13.
  • this airfoil 22 of the compressor stator vane 14 is sep ⁇ arated virtually in a number of airfoil areas, which follows successively one after one.
  • the compressor stator vane 14 comprises again three airfoil sections: a root-sided airfoil area 22a, an in- termediate airfoil area 22b and a tip-sided airfoil area 22c.
  • the respective airfoil areas 22a, 22b, 22c are separated by a first virtual separative profile 21 and a second virtual separative profile 23 from one another.
  • the profiles of the root-sided air ⁇ foil area 22a have stagger angles that are turned close in comparison to the stagger angle of the intermediate airfoil area 22b.
  • the profiles com- prise stagger angles that are larger than the stagger angle of the intermediate airfoil area 22b.
  • Figure 5 shows only schematically a graph 48 for the stagger angles for all profiles of the airfoil 22 of a compressor stator vane 14 over to the span direction from 0% height to 100% height at the free ending airfoil tip.
  • compressor stator vanes 14 comprise a tip-sided airfoil area 22c in which the profile with the largest stagger angle of all profiles of the tip-sided airfoil area 22c is located at about 82% height. With increasing distance to said height the stagger angles of the remaining profiles of the tip-sided airfoil area decreases monotonously towards both the blade tip and the second virtual separative profile 23.
  • the differ ⁇ ence between said largest stagger angle and the stagger angle of the second virtual separative profile 23 is larger than 3°, but not more than 12°.
  • the profiles comprise a stagger angle with only a moderate increase in size with de ⁇ creasing distance to the first virtual separating profile 21.
  • the stagger an ⁇ gles of the profiles are rather uniform.
  • the dif ⁇ ference between the largest and smallest stagger angle of the profiles of the intermediate airfoil area 22b is not greater than 3°.
  • the stagger angle increas- es monotonously significantly i.e. about more than 3°, but not more than 12° from the first virtual separative profile 21 towards the profile at 0% height.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne une pale (42) de compresseur, pour compresseur axial (10), qui comprend une emplanture (19) de pale et un profil aérodynamique (20) en porte-à-faux. La pale de compresseur est fixée à ladite emplanture (19) de pale. Le profil aérodynamique (20, 22) comprend un côté d'aspiration et un côté de pression s'étendant dans une direction d'envergure à partir de ladite emplanture (19) de pale jusqu'à une pointe (24, 28) du profil aérodynamique et dans une direction de corde à partir d'un bord d'attaque (25) côté amont jusqu'à un bord de fuite (27) côté aval. Le profil aérodynamique comprend : une zone de profil aérodynamique côté emplanture (20a, 22a), un profil aérodynamique intermédiaire (20b, 22b) et un profil aérodynamique côté pointe (20c, 22c), pour chaque section de profil du profil aérodynamique (20, 22), un angle de décalage peut être déterminé. Afin de fournir un profil aérodynamique et un compresseur ayant un rendement aérodynamique amélioré, il est proposé que le profil comprenant l'angle de décalage le plus grand de tous les profils du profil aérodynamique côté pointe soit situé entre 78 % de hauteur et 86 % de hauteur et qu'il commence à partir du profil ayant l'angle de décalage le plus grand du profil aérodynamique côté pointe, les angles de décalage des profils restants du profil aérodynamique côté pointe (20c, 22c) vers le second profil virtuel de séparation (23) et vers la pointe de pale diminuant de manière monotone.
PCT/EP2018/069072 2017-07-14 2018-07-13 Pale de compresseur et compresseur axial comprenant ladite pale WO2019012102A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201762532399P 2017-07-14 2017-07-14
US62/532,399 2017-07-14

Publications (1)

Publication Number Publication Date
WO2019012102A1 true WO2019012102A1 (fr) 2019-01-17

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Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2021121456A1 (fr) * 2019-12-20 2021-06-24 MTU Aero Engines AG Aube de turbine à gaz
US11480062B1 (en) 2021-04-30 2022-10-25 General Electric Company Compressor stator vane airfoils
EP4083380A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube rotorique de compresseur
EP4083381A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083388A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083387A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083386A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
EP4083385A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
EP4083382A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
EP4083383A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083384A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083379A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
US11519273B1 (en) 2021-04-30 2022-12-06 General Electric Company Compressor rotor blade airfoils
US11643932B2 (en) 2021-04-30 2023-05-09 General Electric Company Compressor rotor blade airfoils

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EP1505302A1 (fr) * 2003-08-05 2005-02-09 General Electric Company Aube de compresseur
US20080131272A1 (en) * 2006-11-30 2008-06-05 General Electric Company Advanced booster system
EP2827003A1 (fr) * 2013-07-15 2015-01-21 MTU Aero Engines GmbH Grille de guidage de compresseur à turbines à gaz
WO2015178974A2 (fr) * 2014-02-19 2015-11-26 United Technologies Corporation Surface portante de moteur à turbine à gaz

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Publication number Priority date Publication date Assignee Title
EP1505302A1 (fr) * 2003-08-05 2005-02-09 General Electric Company Aube de compresseur
US20080131272A1 (en) * 2006-11-30 2008-06-05 General Electric Company Advanced booster system
EP2827003A1 (fr) * 2013-07-15 2015-01-21 MTU Aero Engines GmbH Grille de guidage de compresseur à turbines à gaz
WO2015178974A2 (fr) * 2014-02-19 2015-11-26 United Technologies Corporation Surface portante de moteur à turbine à gaz

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2021121456A1 (fr) * 2019-12-20 2021-06-24 MTU Aero Engines AG Aube de turbine à gaz
US11927109B2 (en) 2019-12-20 2024-03-12 MTU Aero Engines AG Gas turbine blade
US20230026899A1 (en) * 2019-12-20 2023-01-26 MTU Aero Engines AG Gas turbine blade
EP4083382A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
EP4083384A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083387A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083386A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
EP4083385A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
EP4083381A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083383A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083388A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Profil aérodynamique d'aube rotorique de compresseur
EP4083379A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube statorique de compresseur
US11519272B2 (en) 2021-04-30 2022-12-06 General Electric Company Compressor rotor blade airfoils
US11519273B1 (en) 2021-04-30 2022-12-06 General Electric Company Compressor rotor blade airfoils
EP4083380A1 (fr) * 2021-04-30 2022-11-02 General Electric Company Surface portante d'aube rotorique de compresseur
US11643932B2 (en) 2021-04-30 2023-05-09 General Electric Company Compressor rotor blade airfoils
US11480062B1 (en) 2021-04-30 2022-10-25 General Electric Company Compressor stator vane airfoils
US12018585B2 (en) 2021-04-30 2024-06-25 Ge Infrastructure Technology Llc Compressor rotor blade airfoils

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