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WO2011026468A2 - Turbomachine et procédé de production d'un revêtement de rodage structuré - Google Patents

Turbomachine et procédé de production d'un revêtement de rodage structuré Download PDF

Info

Publication number
WO2011026468A2
WO2011026468A2 PCT/DE2010/001018 DE2010001018W WO2011026468A2 WO 2011026468 A2 WO2011026468 A2 WO 2011026468A2 DE 2010001018 W DE2010001018 W DE 2010001018W WO 2011026468 A2 WO2011026468 A2 WO 2011026468A2
Authority
WO
WIPO (PCT)
Prior art keywords
profile
structured
blade tip
inlet lining
turbomachine
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.)
Ceased
Application number
PCT/DE2010/001018
Other languages
German (de)
English (en)
Other versions
WO2011026468A3 (fr
Inventor
André Werner
Marius Theune
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 GmbH
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 GmbH filed Critical MTU Aero Engines GmbH
Publication of WO2011026468A2 publication Critical patent/WO2011026468A2/fr
Publication of WO2011026468A3 publication Critical patent/WO2011026468A3/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

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
    • 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
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • 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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/10—Manufacture by removing material
    • 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
    • F05D2240/11—Shroud seal segments

Definitions

  • the invention relates to a turbomachine, in particular a gas turbine, with a housing, which has an inlet lining, and a rotor with a plurality of blades, and a method for producing a structured inlet lining.
  • Rotor of gas turbines are known from the prior art, which have blade tip armor in the region of the blade tip, with which they rub against an inlet lining of the turbomachine to produce a minimum running gap.
  • turbomachines with inlet linings are known, in which the inlet lining has certain structures, for example honeycomb structures, which were introduced into the inlet lining in the production of the inlet lining prior to assembly in the turbomachine.
  • the object of the invention is to provide a turbomachine with an easily structured and fluidically optimized inlet lining and a simple method for producing a structured inlet lining.
  • a turbomachine in particular gas turbine, with a housing which has an inlet lining, and a rotor with a plurality of blades, in which the blade tips have an axially structured profile, which generates a counter profile in the inlet lining and which having in the axial direction of the front end region, in particular at the front end itself, its radially outermost point.
  • the structured profile of the blade tip and the counter-profile of the inlet lining form a kind of labyrinth seal between the blade and the inlet lining.
  • turbomachine in particular a gas turbine, having a housing which has an inlet lining, and a rotor having a plurality of blades, in which the blade tips partially remove the inlet lining, and the inlet lining has an axially structured profile in the axial
  • the inlet lining with a structured profile and its radially outermost point in the axially front end region of the blade tip also creates a fluidically optimized turbomachine.
  • the blade tips have a blade tip armor forming the structured profile.
  • the material properties of the blade tip can be optimized for generating the counter profile and for removing the inlet lining.
  • the structured profile of the blade tip can only be formed by the blade tip armor, which simplifies the production of identical rotors with different blade tip profiles.
  • the blade tip armor is, for example, a continuous layer that extends over the entire blade tip, wherein the radial inner side of the layer preferably has a course corresponding to the outer side.
  • the blade tip armor has substantially the same thickness at each location.
  • a geometric profile section which repeats axially a plurality of times to form the structured profile.
  • the repetition of a profile section enhances the effect as a labyrinth seal and simplifies the production.
  • the profile is axially stepped step-shaped.
  • the radial extent of the blade tip can decrease in the axial direction, preferably decrease monotonically and / or steadily, wherein the radially outermost point is present exclusively in the front end region. This allows a further optimization of the fluidic properties of the turbomachine.
  • the inlet lining has radially differently outwardly extending circumferential grooves forming the profile, wherein the radially outermost radially outwardly extending groove opposite the axially forward end region, in particular the end of the blade tip. Due to the different radial extent of
  • the inlet lining may have a structure independent of the blade.
  • the circumferential grooves extend in the circumferential direction or obliquely thereto.
  • the circumferential grooves can have continuous radially inwardly lying groove bottoms compared to each other in the axial direction. This allows additional fluidic optimization.
  • the radial extent of the blade tip decreases in the axial direction as monotonically as the groove bottoms.
  • the circumferential grooves may each have the same radial depth in the inlet lining even with a structured blade tip.
  • the circumferential grooves form a crenellated profile in the circumferential direction.
  • the inlet lining may have at least at one axial end a radially inwardly extending annular edge which radially overlaps the adjacent axial blade tip end.
  • the ring edge acts as a deflection ring for the oncoming gas, which penetrates in a lesser amount in the running gap.
  • the invention further relates to a method for producing a structured inlet lining in a turbomachine described above, wherein the inlet lining is structured by rubbing the rotor blades with a structured blade tip of the rotor of the turbomachine.
  • the inlet lining therefore does not have to have the final structure or even a structure during installation.
  • the inlet lining can also be structured before the rotor blades are scratched. This allows a superimposed structure of a given profile of the inlet lining with the profile produced by the structured blade tip.
  • FIG. 2 is a perspective view of a segment of an inlet lining of a turbomachine according to the invention
  • FIGS. 4a-c show a further embodiment of an inlet lining of a turbomachine according to the invention.
  • FIG. 5a - c an alternative embodiment of an inlet lining of a turbomachine according to the invention.
  • FIGS. 1 a to 1 f show various embodiments of blades 10, which are attached to a rotor of a turbomachine.
  • the turbomachine has an inlet lining 18 as part of a stationary housing, on which the blades 10 of the rotor can touch with their blade tips 12 (see FIG.
  • the axial direction 20 of the turbomachine is indicated by an arrow and corresponds to the main flow direction of the gas.
  • the blade tips 12 have in all embodiments according to Figures 1 a - 1 f blade tip armor 14, which form in the axial direction of structured profiles 16, that is not straight axially. When rubbed against the inlet lining 18, the structured profiles 16 of the blade tips 12 generate corresponding counter profiles in the inlet lining 18.
  • the structured profiles 16 of the blade tips 12 each have their radially outermost point 22 at the front end region of the blade tips 12 which is in the axial direction 20.
  • this point is not only in the axially front fifth of the blade tip armor 14, but at the outermost point of the same.
  • FIG. 1a shows a first embodiment of the blade 10 with the associated inlet lining 18.
  • a blade tip armor 14 is provided which forms the profile 16 structured in the axial direction.
  • the blade 10 itself (below the armor 14) is not structured on its circumferential surface 21, that is, viewed in the circumferential direction, this surface 21 extends axially straight.
  • the blade tip 12 formed by the blade tip armor 14 has the radially outermost point 22 at the axially forward end of the blade 10.
  • FIG. 1b shows an alternative embodiment of the blade tip 12, the profile 16 corresponding to the profile 16 of the first embodiment in FIG. 1a.
  • the blade tip armor 14 is, as in the other embodiments, a continuous layer which extends over the entire blade tip 12, in which case the radial inner side of the layer has a course corresponding to the outer side.
  • the blade tip armor 14 in this embodiment is the same thickness at each point of the blade tip 12.
  • FIG. 1c shows a blade 10 with a crenellated profile.
  • a pinnacle forms a geometric profile section, which repeats axially several times to form the structured profile 16.
  • the pinnacle profile with the axially repeated profile section forms a good labyrinth seal between the blade 10 and inlet lining 18 in the axial direction 20th
  • FIG. 1d Another embodiment with an axially repeated repetitive geometric profile section is shown in Figure 1d, in which the profile 16 is formed like a sawtooth, wherein the teeth extend radially inwardly in the axial direction, to then increase radially straight.
  • Figures 1e and 1f show a profile 16, which is seen in the circumferential direction axially staircase-shaped.
  • the embodiment shown in Figure 1e has two stages and the embodiment shown in Figure 1f has three stages. Again, the steps fall in the axial direction.
  • the blade 10 without blade tip armor 14 has a corresponding profile 16, as shown or claimed.
  • the axially structured counter profile in the inlet lining 18 is, as stated, generated by the brushing of the rotor blades 10 with a structured blade tip 12 during operation of the turbomachine.
  • This allows a simple design of a rotationally symmetrical profile of the inlet lining 18.
  • the fluidic properties of the turbomachine are improved in this way, as on the one hand by the structured profile 16 of the blade 10 and the corresponding generated counter-profile in the inlet lining 18 a kind of labyrinth seal between the blade 10 and inlet lining 18 forms, and on the other hand by selecting the geometry of the profile 16 of the blade tip 12, an aerodynamically optimized inlet lining 18 is formed.
  • a further embodiment provides an inlet lining 18, which is already structured before the rotor blades 10 are scratched, and which is shown in FIGS. 2 to 5.
  • Figure 2 shows in advance a perspective view of a segment of the inlet lining 18 with circumferentially extending circumferential grooves 24, which form a crenellated profile seen in the circumferential direction.
  • FIG. 3 a is a plan view of the inlet lining 18 from FIG. 2.
  • FIGS. 3b and 3c show a sectional view of the inlet lining 18 before or after the coating of a structured blade tip 12.
  • the blade tip 12 shown in FIG. 3b corresponds to the blade tip 12 of the blades 10 from FIGS. 1a and 1b.
  • the circumferential grooves 24 extend radially outwards to different extents, as can be clearly seen in FIG. 3b. Namely, the groove bottoms 26 are continuously radially inward in comparison to one another in the axial direction 20.
  • FIGS. 4a to 4c show an inlet lining 18 in which the circumferential grooves 24 extend at an oblique angle to the axial direction 20.
  • the circumferential grooves 24 extend with increasing axial direction 20 in the direction of rotation 28 of the rotor
  • a radially inwardly extending annular edge 30 is formed, which radially overlaps the adjacent axial blade tip end (see FIG. 4b and FIG. 4c).
  • This overlapping ring edge 30 is also present in FIG. 3c.
  • FIGS. 5a, 5b and 5c show a third embodiment of a structured inlet lining 18.
  • the inlet lining 18 is formed analogously to FIGS. 4a to 4c, wherein the circumferential grooves 24 extend in the axial direction 20 counter to the direction of rotation 28 of the rotor.
  • a structured inlet lining 18 with a structured blade tip 12 allows formation of a complex geometry in the structure of the inlet lining 18, by means of which the turbomachine can be aerodynamically optimized.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une turbomachine, notamment une turbine à gaz, comprenant un logement qui présente un revêtement de rodage (18), un rotor avec plusieurs pales (10), et les pointes de pale (12) présentent un profil (16) structuré dans la direction axiale (20) qui génère un contre-profil dans le revêtement de rodage (18) et présente dans la région d'extrémité avant dans la direction axiale (20), notamment à l'extrémité avant même, son point radial (22) le plus à l'extérieur.
PCT/DE2010/001018 2009-09-04 2010-08-31 Turbomachine et procédé de production d'un revêtement de rodage structuré Ceased WO2011026468A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE200910040298 DE102009040298A1 (de) 2009-09-04 2009-09-04 Strömungsmaschine und Verfahren zur Erzeugung eines strukturierten Einlaufbelags
DE102009040298.5 2009-09-04

Publications (2)

Publication Number Publication Date
WO2011026468A2 true WO2011026468A2 (fr) 2011-03-10
WO2011026468A3 WO2011026468A3 (fr) 2011-10-13

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/DE2010/001018 Ceased WO2011026468A2 (fr) 2009-09-04 2010-08-31 Turbomachine et procédé de production d'un revêtement de rodage structuré

Country Status (2)

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DE (1) DE102009040298A1 (fr)
WO (1) WO2011026468A2 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2012025357A1 (fr) * 2010-08-23 2012-03-01 Rolls-Royce Plc Aube et soufflante correspondante
WO2012110865A1 (fr) * 2011-02-16 2012-08-23 Toyota Jidosha Kabushiki Kaisha Machine rotative
JP2018003841A (ja) * 2016-07-06 2018-01-11 ゼネラル・エレクトリック・カンパニイ タービンロータブレード用シュラウド構成
JP2019100204A (ja) * 2017-11-29 2019-06-24 三菱重工業株式会社 タービン、動翼
US20210189884A1 (en) * 2019-12-20 2021-06-24 United Technologies Corporation Turbine engine rotor blade with castellated tip surface

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102014212652A1 (de) 2014-06-30 2016-01-14 MTU Aero Engines AG Strömungsmaschine

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FR1218301A (fr) * 1958-03-07 1960-05-10 Maschf Augsburg Nuernberg Ag Amélioration de l'étanchéité du joint des aubages mobiles de turbo-machines
US4738586A (en) * 1985-03-11 1988-04-19 United Technologies Corporation Compressor blade tip seal
US4884820A (en) * 1987-05-19 1989-12-05 Union Carbide Corporation Wear resistant, abrasive laser-engraved ceramic or metallic carbide surfaces for rotary labyrinth seal members
EP0661415A1 (fr) * 1993-12-17 1995-07-05 Sulzer Innotec Ag Joint d'étanchéité entre un carter et un corps rotatif
DE4432998C1 (de) * 1994-09-16 1996-04-04 Mtu Muenchen Gmbh Anstreifbelag für metallische Triebwerkskomponente und Herstellungsverfahren
DE10140742B4 (de) * 2000-12-16 2015-02-12 Alstom Technology Ltd. Vorrichtung zur Dichtspaltreduzierung zwischen einer rotierenden und einer stationären Komponente innerhalb einer axial durchströmten Strömungsmaschine
US6568909B2 (en) * 2001-09-26 2003-05-27 General Electric Company Methods and apparatus for improving engine operation
EP1985805B1 (fr) * 2007-04-26 2011-04-06 Siemens Aktiengesellschaft Machine rotative

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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2012025357A1 (fr) * 2010-08-23 2012-03-01 Rolls-Royce Plc Aube et soufflante correspondante
WO2012110865A1 (fr) * 2011-02-16 2012-08-23 Toyota Jidosha Kabushiki Kaisha Machine rotative
US9534503B2 (en) 2011-02-16 2017-01-03 Toyota Jidosha Kabushiki Kaisha Rotary machine
JP2018003841A (ja) * 2016-07-06 2018-01-11 ゼネラル・エレクトリック・カンパニイ タービンロータブレード用シュラウド構成
JP2019100204A (ja) * 2017-11-29 2019-06-24 三菱重工業株式会社 タービン、動翼
US20210189884A1 (en) * 2019-12-20 2021-06-24 United Technologies Corporation Turbine engine rotor blade with castellated tip surface
US11225874B2 (en) * 2019-12-20 2022-01-18 Raytheon Technologies Corporation Turbine engine rotor blade with castellated tip surface

Also Published As

Publication number Publication date
DE102009040298A1 (de) 2011-03-10
WO2011026468A3 (fr) 2011-10-13

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