EP2696042A1 - Turbomachine avec au moins un stator - Google Patents

Turbomachine avec au moins un stator Download PDF

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Publication number
EP2696042A1
EP2696042A1 EP12179779.9A EP12179779A EP2696042A1 EP 2696042 A1 EP2696042 A1 EP 2696042A1 EP 12179779 A EP12179779 A EP 12179779A EP 2696042 A1 EP2696042 A1 EP 2696042A1
Authority
EP
European Patent Office
Prior art keywords
vane
taper
vanes
turbomachine
guide vanes
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.)
Granted
Application number
EP12179779.9A
Other languages
German (de)
English (en)
Other versions
EP2696042B1 (fr
Inventor
Roland Dr. Wunderer
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines 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 MTU Aero Engines AG filed Critical MTU Aero Engines AG
Priority to EP12179779.9A priority Critical patent/EP2696042B1/fr
Priority to US13/960,943 priority patent/US9506360B2/en
Publication of EP2696042A1 publication Critical patent/EP2696042A1/fr
Application granted granted Critical
Publication of EP2696042B1 publication Critical patent/EP2696042B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • 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
    • F01D5/146—Shape, i.e. outer, aerodynamic form of blades with tandem configuration, split blades or slotted 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • 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
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
    • 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
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge 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
    • 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
    • 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
    • F05D2270/102—Compressor surge or stall caused by working fluid flow velocity profile distortion

Definitions

  • the invention relates to a turbomachine, in particular an axial compressor, with at least one vane ring and a method for increasing the stability of the flow in a turbomachine.
  • adjustable vanes are often present, especially at the front stages of compressors.
  • the inflow angles to the following blades are thus set and the energy conversion of the stage formed by the vanes and the rotor blades is controlled.
  • the vanes are adjusted, the flow angle changes over the entire channel height.
  • the distribution of the local mass flow changes along the channel height. As a result, the stability of the flow in the turbomachine can be reduced and the efficiency can be reduced.
  • the EP 0 745 755 A1 a specially shaped guide vane for a compressor of a gas turbine.
  • the vane has an angled towards the blade root rear edge.
  • the use of such vanes improves the stability of the flow and thus increases the compressor pumping limit.
  • the disadvantage here is that the geometry of the vanes is adapted to a particular operating condition at a deviation from the operating condition no improved flow is more guaranteed.
  • the DE 10 2009 023 100 A1 describes a vane device with vanes arranged downstream of each other in the flow direction, wherein the trailing edges of the upstream vanes are shaped differently than the leading edges of the downstream vane, resulting in an uneven distance along the vane edges. Also by this arrangement, the flow in the turbomachine is to be stabilized. The arrangement also has the Disadvantage on that the geometry of the blades is adapted to a specific operating condition.
  • the WO 2007/042522 A1 describes a blade for a turbomachine in which the chord length is uneven along the blade length. This bucket minimizes blade lattice losses and is also designed for a specific operating range.
  • the invention is therefore based on the object to provide a turbomachine in which the stability of the flow improves and the operating range is extended.
  • the object is achieved in a turbomachine, in particular an axial compressor, with at least one vane ring comprising at least one row of adjustable vanes, wherein each vane in a side view of the vane with respect to its blade in the direction of its longitudinal axis has a taper.
  • Each row of vanes includes first vanes and second vanes, wherein in a common side view of a first vane and a second vane, each first vane has a taper in a longitudinal direction along its airfoil and every other vane has a taper in an opposite direction.
  • the common side view is created by a juxtaposition of a loose first vane and a loose second vane.
  • the common side view is not the view of a first vane and a second vane in their installed position.
  • each vane affects each associated airfoil.
  • the counter-tapering of the vanes causes the flow to be more deflected in the unrestrained areas than in the tapered areas. This stabilizes the flow.
  • a preferred field of application are axial compressors.
  • At least a first vane and at least one second vane form a unit, and a plurality of these units are uniformly distributed in the circumferential direction of the vane ring, wherein the distances between adjacent vanes different or equal are big.
  • the second guide vanes are arranged opposite the first guide vanes of the same row in the axial direction of the turbomachine in the same position or offset from one another. At the same position results in a compact design.
  • the flow guidance can be extended in the axial direction of the turbomachine.
  • the first guide vanes and the two guide vanes may have a covering region, wherein the covering region has a straight, oblique or curved course in the axial direction of the turbomachine.
  • the overlap area allows in the radial direction of the turbomachine a flowing transition between the inner and outer flow guide in the channels between two adjacent vanes.
  • the coverage area may range from 30% to 70% of a channel height defined by a length of the airfoil of each first vane and each second vane. This area gives the best results.
  • the taper of each first vane at the trailing edge of the respective first vane and the taper of each second vane at the trailing edge of the respective second vane are formed.
  • the taper of each first vane at the leading edge of the respective first vane and the taper of each second vane at the leading edge of the respective second vane are formed.
  • the taper of each first vane at the trailing edge of the respective first vane and the taper of each second vane at the leading edge of the respective second vane are formed.
  • the taper of each first vane at the trailing edge and the leading edge of the respective first vane and the taper of each second vane are also formed at the trailing edge and the leading edge of the respective second vane.
  • the taper can each be formed by a single or double curvature or the taper can be formed in each case by a gradation with at least two step transitions, each step transition is edged or rounded.
  • a curvature allows a flowing deflection of the flow, while a gradation is easier to produce.
  • each first vane may be 30% to 70% of the maximum width of each second vane, or the taper of each second vane may be 30% to 70% of the maximum width of each first vane. This area gives the best results.
  • first vanes and the second vanes may each have different curvatures of the blade skeleton lines and / or the different profiles in the same position along the longitudinal axis. This allows a further, more detailed adaptation of the channels to the local flow.
  • the first guide vanes and the second guide vanes may each have different curvatures of the skeleton lines and / or different profiles in an outer region and an inner region.
  • the overlap area ensures a smooth transition between the areas lying in the radial direction of the turbomachine outside and inside.
  • first vanes and the second vanes are each rotatable about their longitudinal axis, wherein the first vanes and the second vanes are coupled or independently rotatable.
  • the guide vane grille can be adapted to the operating state. The separate adjustment of the first and second vanes allows even more specific adaptation of the ducts.
  • each internally-tapered vane deflects the flow outwardly and each outwardly-tapered vane deflects the flow internally.
  • the increased stability of the flow prevents compressor pumping and allows a larger operating range of the turbomachine.
  • the Fig. 1 to 5 each show a schematic side view in common of a first vane 1 and a second vane 2 of a vane ring, not shown, of a turbomachine also not shown.
  • the blade roots and the blade heads of the first vanes 1 and the second vanes 2 are not shown, so that the illustration refers to the airfoils.
  • the first vane 1 is shown in each case with a solid line.
  • Each first vane 1 has a leading edge 1a, a trailing edge 1b, and a taper 1c.
  • the second vane 2 is shown in each case with a dashed line.
  • Each second vane 2 has a leading edge 2a, a trailing edge 2b, and a taper 2c.
  • first vanes 1 and the second vanes 2 have a common longitudinal axis 3 which is at the same time an axis of rotation.
  • the longitudinal axes or axes of rotation of the first guide vanes 1 and the second guide vanes 2 can also be located at different positions in the axial direction of the turbomachine.
  • the first vanes 1 and second vanes 2 illustrated with respect to the airfoils define a channel height 4 having an inner portion 4a and an outer portion 4b in the radial direction of the turbomachine, not shown. Between the inner area 4a and the outer area 4b there is a covering area 4c.
  • Fig. 1 is the taper 1c of the first vane 1 at the trailing edge 1b in the outer region 4b.
  • the taper 2c of the second vane 2 is located at the trailing edge 2b in the inner region 4a.
  • the covering area 4c has a straight course.
  • Fig. 2 is the taper 1c of the first vane 1 at the front edge 1a in the inner region 4a.
  • the taper 2c of the second vane 2 is located at the leading edge 2a in the outer region 4b.
  • the covering area 4c has a straight course.
  • Fig. 3 is the taper 1c of the first vane 1 at the trailing edge 1b in the inner region 4a.
  • the taper 2c of the second vane 2 is located at the leading edge 2a in the outer region 4b.
  • the covering area 4c has a straight course.
  • Fig. 4 is the taper 1c of the first vane 1 at the trailing edge 1b in the outer region 4b.
  • the taper 2c of the second vane 2 is located at the leading edge 2a in the inner region 4a.
  • the covering area 4c has a straight course.
  • Fig. 5 is the taper 1c of the first vane 1 at the leading edge 1a and at the trailing edge 1 b in the inner region 4a.
  • the taper 2c of the second vane 2 is located at the leading edge 2a and at the trailing edge 2b in the outer region 4b.
  • the covering area 4c has a curved course.
  • the flow at the taper 1c of each first vane 1 and at the taper 2c of each second vane 2 is locally less deflected than where there is no taper 1c on each first vane 1 and no taper 2c on each second vane 2.
  • a leading edge 1a, 2a or a trailing edge 1b, 2b of a first vane 1 and a second vane 2 are shown in the overlapping area 4c.
  • the front edge 1a, 2a or the trailing edge 1b, 2b has two step transitions 5.
  • the step transitions 5 are formed at right angles.
  • Fig. 6b The front edge 1a, 2a or the trailing edge 1b, 2b has two step transitions 5. Between the step transitions 5 is a slope. 6
  • first vanes 1 and second vanes 2 are shown in a row of a vane ring, not shown here.
  • the arrangement is formed in each case from units 8 in which a specific combination of first guide vanes 1 and second guide vanes 2 is fixed.
  • the units 8 are arranged uniformly along the entire inner periphery of the turbomachine, not shown.
  • the unit 8 comprises alternately arranged first guide vanes 1 and second guide vanes 2, which occupy the same position in the axial direction of the turbomachine, not shown.
  • the unit 8 comprises alternately arranged first guide vanes 1 and second guide vanes 2, which are arranged offset in the axial direction of the turbomachine, not shown.
  • the unit 8 comprises two adjacent guide vanes 1 and a second guide blade 2, which are arranged offset in the axial direction of the turbomachine, not shown.
  • first guide vanes 1 with second guide vanes 2 in the ratio of, for example, 3: 1 or 1: 1 is possible.
  • a ratio of 2: 1 is preferred.
  • Fig. 10 sketched is also a unit 8 with a "mirrored" combination possible. So is for example in FIG. 10 a combination of a first vane 1 with two second vanes 2 shown. Likewise, combination of first guide vanes 1 with second guide vanes 2 in the ratio of, for example, 1: 3 is possible.
  • FIG. 11 another unit 8 is shown. In this embodiment, there is no axial overlap between the first blades 1 and the second blades 2.
  • Fig. 12 to 14 In each case, a perspective view of an outwardly open guide vane ring 9 of the turbomachine, not shown, with first guide vanes 1 and second guide vanes 2 is shown.
  • Fig. 12 The first vanes 1 and the second vanes 2 correspond to the Fig. 1 . 6c and 7 ,
  • Fig. 13 The first vanes 1 and the second vanes 2 correspond to the Fig. 4 . 6c and 7 ,
  • Fig. 14 The first vanes 1 and the second vanes 2 correspond to the Fig. 4 . 6c and 8th ,
  • Turbomachine in particular axial compressor, with at least one vane ring, which comprises at least one row of adjustable guide vanes, wherein each vane in a side view of the vane with respect to its blade in the direction of its longitudinal axis has a taper.
  • each row of vanes includes first vanes and second vanes, wherein in a common side view of a first vane and a second vane, each first vane has a taper in a longitudinal direction along its airfoil and every other vane having a taper in an opposite direction.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP12179779.9A 2012-08-09 2012-08-09 Turbomachine avec au moins un stator Active EP2696042B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP12179779.9A EP2696042B1 (fr) 2012-08-09 2012-08-09 Turbomachine avec au moins un stator
US13/960,943 US9506360B2 (en) 2012-08-09 2013-08-07 Continuous-flow machine with at least one guide vane ring

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP12179779.9A EP2696042B1 (fr) 2012-08-09 2012-08-09 Turbomachine avec au moins un stator

Publications (2)

Publication Number Publication Date
EP2696042A1 true EP2696042A1 (fr) 2014-02-12
EP2696042B1 EP2696042B1 (fr) 2015-01-21

Family

ID=46642433

Family Applications (1)

Application Number Title Priority Date Filing Date
EP12179779.9A Active EP2696042B1 (fr) 2012-08-09 2012-08-09 Turbomachine avec au moins un stator

Country Status (2)

Country Link
US (1) US9506360B2 (fr)
EP (1) EP2696042B1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3686396A1 (fr) * 2019-01-24 2020-07-29 MTU Aero Engines GmbH Grille d'aube de guidage pour une turbomachine

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Publication number Priority date Publication date Assignee Title
US20160146040A1 (en) * 2014-11-25 2016-05-26 United Technologies Corporation Alternating Vane Asymmetry
TWI678471B (zh) * 2018-08-02 2019-12-01 宏碁股份有限公司 散熱風扇
US11401824B2 (en) * 2019-10-15 2022-08-02 General Electric Company Gas turbine engine outlet guide vane assembly
US11686321B2 (en) * 2021-11-10 2023-06-27 Air Cool Industrial Co., Ltd. Ceiling fan having double-layer blades

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US4995786A (en) * 1989-09-28 1991-02-26 United Technologies Corporation Dual variable camber compressor stator vane
EP1998006A2 (fr) * 2007-05-31 2008-12-03 United Technologies Corporation Mécanisme de rétention pour aube de guidage d'admission
WO2010007224A1 (fr) * 2008-06-25 2010-01-21 Snecma Compresseur de turbomachine

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US6375419B1 (en) * 1995-06-02 2002-04-23 United Technologies Corporation Flow directing element for a turbine engine
US6328533B1 (en) * 1999-12-21 2001-12-11 General Electric Company Swept barrel airfoil
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JP4786077B2 (ja) 2001-08-10 2011-10-05 本田技研工業株式会社 タービン用静翼及びその製造方法
US6554564B1 (en) 2001-11-14 2003-04-29 United Technologies Corporation Reduced noise fan exit guide vane configuration for turbofan engines
SE527786C2 (sv) * 2004-11-05 2006-06-07 Volvo Aero Corp Stator till en jetmotor och en jetmotor innefattande sådan stator
DE112006002658B4 (de) * 2005-10-11 2021-01-07 General Electric Technology Gmbh Turbomaschinenschaufel
EP2140111B1 (fr) * 2007-04-24 2014-05-07 Alstom Technology Ltd Turbomachine
EP2199544B1 (fr) * 2008-12-22 2016-03-30 Techspace Aero S.A. Architecture de redresseur
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Publication number Priority date Publication date Assignee Title
US4995786A (en) * 1989-09-28 1991-02-26 United Technologies Corporation Dual variable camber compressor stator vane
EP1998006A2 (fr) * 2007-05-31 2008-12-03 United Technologies Corporation Mécanisme de rétention pour aube de guidage d'admission
WO2010007224A1 (fr) * 2008-06-25 2010-01-21 Snecma Compresseur de turbomachine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3686396A1 (fr) * 2019-01-24 2020-07-29 MTU Aero Engines GmbH Grille d'aube de guidage pour une turbomachine
US11280212B2 (en) 2019-01-24 2022-03-22 MTU Aero Engines AG Guide vane cascade for a turbomachine

Also Published As

Publication number Publication date
US9506360B2 (en) 2016-11-29
EP2696042B1 (fr) 2015-01-21
US20140044518A1 (en) 2014-02-13

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