EP1832717A1 - Procédé pour modifier le flux d'air de bout d'aube dans une turbomachine axiale et canal annulaire pour l'écoulement axial du fluide dans une turbomachine - Google Patents

Procédé pour modifier le flux d'air de bout d'aube dans une turbomachine axiale et canal annulaire pour l'écoulement axial du fluide dans une turbomachine Download PDF

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Publication number
EP1832717A1
EP1832717A1 EP06004866A EP06004866A EP1832717A1 EP 1832717 A1 EP1832717 A1 EP 1832717A1 EP 06004866 A EP06004866 A EP 06004866A EP 06004866 A EP06004866 A EP 06004866A EP 1832717 A1 EP1832717 A1 EP 1832717A1
Authority
EP
European Patent Office
Prior art keywords
flow
flow channel
partial
main
pressure ratio
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Withdrawn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
EP06004866A
Other languages
German (de)
English (en)
Inventor
Romain Bayere
Malte Dr. Blomeyer
Christian Dr. Cornelius
Torsten Matthias
Uwe Dr. Sieber
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.)
Siemens AG
Original Assignee
Siemens 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 Siemens AG filed Critical Siemens AG
Priority to EP06004866A priority Critical patent/EP1832717A1/fr
Publication of EP1832717A1 publication Critical patent/EP1832717A1/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Withdrawn legal-status Critical Current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
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • 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
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/105—Final actuators by passing part of the fluid
    • 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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/02—Surge control
    • F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
    • 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
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526—Details of the casing section radially opposing blade tips
    • 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
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/685—Inducing localised fluid recirculation in the stator-rotor interface
    • 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
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/10—Purpose of the control system to cope with, or avoid, compressor flow instabilities
    • F05D2270/101—Compressor surge or stall

Definitions

  • the invention relates to a method for influencing the near-gap flow, wherein in each case a gap is provided between the tips of the blades of a blade ring and a boundary wall opposite the tips of a turbomachine, coupled from a mainstream flowing through the flow mainstream at least a partial flow and upstream of its removal position the main flow is fed again.
  • the invention relates to an annular flow channel for a flowable in the axial direction of a main flow machine, which is concentrically disposed about an axially extending central axis and which is bounded by a circular cross-section boundary wall, wherein an axial wall portion of the boundary wall, which at least the tips of Flow channel radially profiled blades of a blade ring is in each case opposite gap formation, a plurality of distributed over the circumference return flow regions through each of which a main stream at a removal position auskoppelbarer partial flow at an upstream in relation to feed position in the main stream is traceable.
  • Gas turbines and their functions are well known.
  • the sucked by a compressor of the gas turbine air is compressed in this and then mixed with fuel in a burner.
  • the subsequently flowing into a combustion chamber mixture burns to a hot gas, which then flows through a turbine downstream of the turbine and meanwhile offset due to its relaxation, the rotor of the gas turbine in rotation.
  • Both the compressor and the turbine each consist of several successively connected blade stages, each comprising two successive wreaths of blades.
  • a turbine stage is composed of a stator vane formed by non-rotating vanes and a rotor blade disposed downstream thereof, whereas a stage disposed in the compressor is composed of a rotor blade and a stator vane arranged downstream thereof; each viewed in the flow direction of the medium flowing through.
  • all blades are permanently mounted on the common rotor.
  • the series-arranged, i. axially successive compressor stages promote due to the rotating rotor blades with the sucked air from the inlet of the compressor in the direction of the compressor outlet, wherein the air within each stage (or ring) undergoes an incremental pressure increase.
  • the total pressure increase across the compressor is the sum of all incremental pressure rises across each stage (or all of the rings).
  • a stall occurs on one or more airfoils within the compressor, where the flow of air in the main flow direction through part of a compressor stage ceases the energy transferred from the rotor to the air is insufficient to convey the air through the compressor stage and to produce the required pressure ratio of the compressor stage concerned.
  • the pressure ratio is the pressure increase occurring over the relevant stage of the compressor, based on the inlet pressure of the respective stage. If the stall is not immediately counteracted, its progression may be cause all the air flow through the compressor to reverse direction, known as compressor pumps. This particularly critical operating condition endangers the blading and prevents sufficient supply of the combustion chamber with compressor air, so that a faulty operation of the gas turbine must be diagnosed.
  • EP 1 286 022 A1 A similar device and a similar method are known from EP 1 286 022 A1 known.
  • a method for influencing the near-gap flow and a suitable annular flow channel for an axial flow-through turbomachine, wherein the mass flow of the decoupled partial flow is adjusted by means of an adjusting member.
  • the device according to the invention in at least one of the return flow areas on an adjusting member for adjusting the mass flow of the decoupled partial flow.
  • the compressor or the compressor stage or the blade ring is then operated in the stationary operating state, which allows a continuous, reliable and efficient delivery of the medium to be compressed in the downstream direction.
  • the partial flow decoupled from the main flow is taken off in the region of the trailing edges of the blades of the blade ring.
  • the inlet opening for decoupling of the partial flow in the remindström Scheme is completely or partially downstream of the trailing edge of the opposite Airfoil, since the partial flow driving the pressure of the main flow at this point is relatively constant. This leads to a particularly continuous partial flow, which has a particularly positive influence on the gap or wall-near flow.
  • the partial wall flow taken from the boundary wall side is supplied to the main flow in the area of the leading edges of the blades of the blade ring.
  • the outflow opening of the partial flow is completely or partially directly upstream of the leading edge of the opposite airfoil. This results locally locally improved flow conditions in the area of the leading edges of the blades, as a result of which the pressure ratio to be overcome in this area, ie at the tips of the blades, is lowered slightly, so that a tearing off of the flow in the blade tip area can be effectively avoided.
  • the method in which the mass flow of the partial flow is controlled or regulated.
  • the desired or required mass flow can be adjusted.
  • Either all blades of the stage and / or the blade ring are operated in the optimum operating range without overflow, ie there is no partial flow, or, at elevated Likelihood of stall or compressor-pumping, it can be positively influenced the gap near flow by means of a flowing partial flow, with overflow so.
  • connection of the partial flow leads to a pulse in the flow of the airfoil, which also counteracts the impending stall.
  • the adjustment or regulation of the mass flow of the partial flow takes place as a function of the actually occurring pressure ratio of the blade ring, which pressure ratio indicates the downstream edge pressure of the main flow with respect to the inflow edge side pressure.
  • the regulation takes place in such a way that, if the actually occurring pressure ratio is below a predetermined limiting pressure ratio, the flow of a partial flow is prevented. Only when the actually occurring pressure ratio is above the limiting pressure ratio, a partial flow flows.
  • the regulation may even be such that, as the distance between the actually occurring pressure ratio and the limiting pressure ratio decreases, the mass flow of the partial flow is increased linearly or not linearly.
  • the limiting pressure ratio is as close to the maximum, based on the blades of the blade ring pressure ratio as possible or corresponds to this.
  • each return flow region has an adjusting member, so that evenly over the circumference of the annular flow channel a uniform influencing of the gap and wall near flow can be set for each return flow region.
  • the return flow areas are each formed as a return flow channel which extends in the interior of the boundary wall in the axial direction.
  • the arranged in the wall portion return channel then has an inflow opening for the partial flow, the axial position is partially or preferably completely downstream of the trailing edge of the opposite airfoil.
  • the outflow opening of the return flow channel arranged in the wall section then lies in the region of the leading edge of the blade profile, that is to say partially or preferably completely upstream thereof.
  • the adjusting member or each adjustment member which may be formed as a slide, throttle or valve, connected to a control or a control device for adjusting the mass flow to adjust these preferably depending on the respective blade ring outlet pressure.
  • the flow channel has a measuring device, which, with respect to the blade, has downstream pressure of the main flow and which is connected to the adjusting member is in communication via the control or regulating device.
  • FIG. 1 shows a gas turbine 1 in a longitudinal partial section. It has inside a rotatably mounted about a rotation axis 2 rotor 3, which is also referred to as a turbine runner. Along the rotor 3 successive an intake 4, a compressor 5, a toroidal annular combustion chamber 6 with a plurality of rotationally symmetrical to each other arranged burners 7, a turbine unit 8 and an exhaust housing 9.
  • the annular combustion chamber 6 forms a combustion chamber 17 which communicates with an annular hot gas channel 18.
  • the turbine unit 8 is connected in series with one another. Each turbine stage 10 is formed from two blade rings.
  • vanes 13 formed by a blade 15 row 14.
  • the vanes 12 are attached to the stator, whereas the blades 15 a row 14 each by means of a turbine disk 19 on the rotor 3 are attached.
  • a generator or a working machine (not shown) is coupled.
  • FIG. 2 shows in a cross section the input-side section of the compressor 5.
  • the main flow direction 21 coincides with the axial direction.
  • the arranged on the rotor 3 inner boundary wall 27 is formed by the platforms of the rotor mounted in a ring 13 arranged blades 15.
  • Each blade 15 has an aerodynamically optimized profiled blade 22, the tips 33 each have a gap 35 forming the outer boundary wall 25 opposite.
  • a vane 12 of the vane ring 14 fixed to the outer shell side boundary wall 25 instead of this guide vane 12 may be provided at this point also with respect to the axis of rotation 2 of the rotor 3 about the radial direction rotatable inlet guide vane.
  • a vane 12 of the blade ring 14 is also shown downstream of the blade 15, which is also attached to the housing or on the outer boundary wall 25.
  • These two blade rings 13, 14 form a compressor stage.
  • Each vane 12 closes with its opposite end of the rotor 3 also with the boundary wall 27 a gap.
  • a casing treatment can be provided on the boundary wall 27 and a casing treatment can be provided.
  • each return flow channel 41 has at an extraction position 45 an inflow opening for decoupling a partial flow 49 from the main flow 32.
  • Each return flow channel 41 opens into an outflow opening located at a feed position 47, always in relation to the main flow direction 31, upstream of the inflow opening.
  • a partial flow 49 can be flowed, which is in the flow channel 29 flowing main flow 32, such as air, coupled out and this upstream of the removal position 45 is traceable again.
  • an adjusting member 51 is provided for adjusting the partial flow 49 flowing through the return flow region 43.
  • the adjusting member 51 may be formed as a slide 52, throttle or valve (FIG 4) and are in communication with a control or regulating device, so that they can act on the adjusting member 51 to adjust the mass flow of the partial flow 49.
  • the removal position 45 for the disengageable partial flow 49 is axially seen in the wall portion of the boundary wall 25, which downstream of the trailing edge 53 of the opposing profiled airfoil 22, based on the flowing in the main flow direction 31 the annular flow channel 29 medium 30. Accordingly, under the wall section also to understand the immediately and slightly further downstream of the trailing edge 53 lying region of the boundary wall 25.
  • the axial feed position 47 of the return flow region 43, ie, the mouth of the return flow channel 41, is located in the region of the leading edge 55 of the wall section opposite blades 22, preferably completely immediately upstream thereof. Consequently, this area is also to be understood as meaning the axial wall section of the boundary wall 25, which lies directly upstream of the leading edges 55 of the blade 22.
  • FIG. 3 shows, in abstracted form, a plan view (in the radial direction) of the tip 33 of a blade 15, in which the return flow regions 43 uniformly distributed over the circumference in the outer boundary wall 25 can be seen as return flow channels 41.
  • the boundary wall 25 itself is hidden here.
  • the airfoil 22 is first of all flown against at its leading edge 55 by a medium 30 flowing in the main flow direction 31.
  • the coinciding with the circumferential direction of rotation of the rotor 3 rotating blade 15 is indicated by the arrow U.
  • a common slide 52 is provided therein as an adjustment member 51 by means of which the minimum cross section of each return flow region 43 can be varied.
  • the example comb-like slide 52 is displaceable in the circumferential direction U and each has a projecting into each return flow channel 41 tooth, which is also pushed out to set the minimum cross-section of this.
  • valve 51 is provided as adjusting.
  • At least one downstream pressure of the main flow 32 detecting measuring device is provided, which is indirectly connected to the adjusting member 51 via a control or regulating device, so that they can cause the need-based adjustment.
  • the invention provides a flow channel 29 for a through-flow of a main flow 32 in the axial flow machine, wherein the flow channel 29 is concentrically disposed about an axially extending central axis 2 and which is bounded by a circular cross-section boundary wall 25, 27, wherein a axial wall portion of the boundary wall 25, 27, which approximately opposite the tips 33 of the flow channel 29 radially arranged profiled blades 22 of a blade ring 13, 14 in each case with gap formation.
  • return flow areas 41 are provided, through which a return flow 49, which can be removed from the main flow 32, is decoupled and can be supplied again upstream of the coupling-out position.
  • the return flow regions 41 are equipped with adjusting members 51. Furthermore, a method for influencing the gap-near flow is specified, whereby the risk of flow separation on the blade blades is significantly reduced.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Sustainable Development (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP06004866A 2006-03-09 2006-03-09 Procédé pour modifier le flux d'air de bout d'aube dans une turbomachine axiale et canal annulaire pour l'écoulement axial du fluide dans une turbomachine Withdrawn EP1832717A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP06004866A EP1832717A1 (fr) 2006-03-09 2006-03-09 Procédé pour modifier le flux d'air de bout d'aube dans une turbomachine axiale et canal annulaire pour l'écoulement axial du fluide dans une turbomachine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP06004866A EP1832717A1 (fr) 2006-03-09 2006-03-09 Procédé pour modifier le flux d'air de bout d'aube dans une turbomachine axiale et canal annulaire pour l'écoulement axial du fluide dans une turbomachine

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Publication Number Publication Date
EP1832717A1 true EP1832717A1 (fr) 2007-09-12

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EP06004866A Withdrawn EP1832717A1 (fr) 2006-03-09 2006-03-09 Procédé pour modifier le flux d'air de bout d'aube dans une turbomachine axiale et canal annulaire pour l'écoulement axial du fluide dans une turbomachine

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EP (1) EP1832717A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104196746A (zh) * 2014-08-04 2014-12-10 河北瑞兆激光再制造技术有限公司 轴流风机修复后径向间隙控制检验平台的研制方法
EP3179113A1 (fr) * 2015-12-08 2017-06-14 General Electric Company Traitement de mur d'extrémité à effet venturi
RU2715459C1 (ru) * 2019-06-07 2020-02-28 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения имени П.И. Баранова" Турбокомпрессор с надроторным устройством
US10876549B2 (en) 2019-04-05 2020-12-29 Pratt & Whitney Canada Corp. Tandem stators with flow recirculation conduit
US11441575B2 (en) * 2020-02-26 2022-09-13 Honda Motor Co., Ltd. Axial compressor

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE881743C (de) * 1944-01-06 1953-07-02 Messerschmitt Boelkow Blohm Verfahren zur Verhuetung des Abreissens der Stroemung in Verdichtern von Heissstrahltriebwerken
US3572960A (en) * 1969-01-02 1971-03-30 Gen Electric Reduction of sound in gas turbine engines
US4580943A (en) * 1980-12-29 1986-04-08 The United States Of America As Represented By The Secretary Of The Army Turbine wheel for hot gas turbine engine
EP0606475A1 (fr) * 1991-10-04 1994-07-20 Ebara Corporation Turbomachine
JPH06207558A (ja) * 1993-01-11 1994-07-26 Ishikawajima Harima Heavy Ind Co Ltd エンジン用ファンの作動安定化装置
US5431533A (en) * 1993-10-15 1995-07-11 United Technologies Corporation Active vaned passage casing treatment
EP0719907A1 (fr) 1994-12-29 1996-07-03 United Technologies Corporation Virole pour turbine à gaz
EP1286022A1 (fr) 2001-08-14 2003-02-26 United Technologies Corporation Traitement de l'enveloppe pour compresseurs

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE881743C (de) * 1944-01-06 1953-07-02 Messerschmitt Boelkow Blohm Verfahren zur Verhuetung des Abreissens der Stroemung in Verdichtern von Heissstrahltriebwerken
US3572960A (en) * 1969-01-02 1971-03-30 Gen Electric Reduction of sound in gas turbine engines
US4580943A (en) * 1980-12-29 1986-04-08 The United States Of America As Represented By The Secretary Of The Army Turbine wheel for hot gas turbine engine
EP0606475A1 (fr) * 1991-10-04 1994-07-20 Ebara Corporation Turbomachine
JPH06207558A (ja) * 1993-01-11 1994-07-26 Ishikawajima Harima Heavy Ind Co Ltd エンジン用ファンの作動安定化装置
US5431533A (en) * 1993-10-15 1995-07-11 United Technologies Corporation Active vaned passage casing treatment
EP0719907A1 (fr) 1994-12-29 1996-07-03 United Technologies Corporation Virole pour turbine à gaz
EP1286022A1 (fr) 2001-08-14 2003-02-26 United Technologies Corporation Traitement de l'enveloppe pour compresseurs

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
PATENT ABSTRACTS OF JAPAN vol. 018, no. 569 (M - 1695) 31 October 1994 (1994-10-31) *

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104196746A (zh) * 2014-08-04 2014-12-10 河北瑞兆激光再制造技术有限公司 轴流风机修复后径向间隙控制检验平台的研制方法
CN104196746B (zh) * 2014-08-04 2016-02-03 河北瑞兆激光再制造技术有限公司 轴流风机修复后径向间隙控制检验平台的研制方法
EP3179113A1 (fr) * 2015-12-08 2017-06-14 General Electric Company Traitement de mur d'extrémité à effet venturi
US10041500B2 (en) 2015-12-08 2018-08-07 General Electric Company Venturi effect endwall treatment
US10876549B2 (en) 2019-04-05 2020-12-29 Pratt & Whitney Canada Corp. Tandem stators with flow recirculation conduit
RU2715459C1 (ru) * 2019-06-07 2020-02-28 Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения имени П.И. Баранова" Турбокомпрессор с надроторным устройством
US11441575B2 (en) * 2020-02-26 2022-09-13 Honda Motor Co., Ltd. Axial compressor

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