EP3369892A1 - Contournage d'une plate-forme de grille d'aube - Google Patents

Contournage d'une plate-forme de grille d'aube Download PDF

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Publication number
EP3369892A1
EP3369892A1 EP17159076.3A EP17159076A EP3369892A1 EP 3369892 A1 EP3369892 A1 EP 3369892A1 EP 17159076 A EP17159076 A EP 17159076A EP 3369892 A1 EP3369892 A1 EP 3369892A1
Authority
EP
European Patent Office
Prior art keywords
blade
platform
edge
segment
upstream
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
EP17159076.3A
Other languages
German (de)
English (en)
Other versions
EP3369892B1 (fr
Inventor
Inga Mahle
Markus Brettschneider
Fadi Maatouk
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
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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 EP17159076.3A priority Critical patent/EP3369892B1/fr
Priority to ES17159076T priority patent/ES2819128T3/es
Priority to US15/908,401 priority patent/US10648339B2/en
Publication of EP3369892A1 publication Critical patent/EP3369892A1/fr
Application granted granted Critical
Publication of EP3369892B1 publication Critical patent/EP3369892B1/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
    • 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
    • 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
    • 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
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • 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
    • F05D2240/129—Cascades, i.e. assemblies of similar profiles acting in parallel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades

Definitions

  • the present invention relates to a blade grid segment, a blade grid, a platform and a blade channel of a turbomachine and a turbomachine.
  • Turbomachines such as gas and steam turbines regularly have a flow channel for the passage of a fluid.
  • the flow channel also referred to as an annulus, is bounded radially inwardly by the shaft of a rotor and radially outward by a housing, the terms "radial” as well as “axial” and “circumferential direction” and terms derived therefrom are in this Scripture - unless otherwise stated - always to understand a (intended) rotation axis of the rotor.
  • blade grids are arranged (for which the term “blade ring” is common). They each include vanes or blades that are circumferentially spaced one behind the other at substantially regular intervals, and associated platforms, also referred to as “deck plates", each having an upstream and a downstream platform edge. These platform edges limit the blade surface (or blades) facing platform surface in the axial direction.
  • the edge of the platform referred to as the "upstream” platform edge is referred to as the (axial) main flow passing through the annulus of the turbomachine during operation, as the “downstream edge” corresponding to the other edge.
  • the statements “downstream” and “upstream” respectively refer to the main axial flow direction and thereby only to the axial position, that is to say regardless of a possible displacement in the circumferential direction.
  • a point is to be understood as “downstream of the leading edges”. if it is arranged offset to a direct connection of the leading edges on the platform surface axially in the main flow direction.
  • the pressure side of a blade and the suction side of an adjacent blade define in the circumferential direction in each case a so-called blade channel.
  • a blade channel In the radial direction of this blade channel is limited within the turbomachine by so-called side walls.
  • blade intermediate strip The section of the platform surface, which in the axial direction by the direct connections of the leading edges or the trailing edges of adjacent blades on the platform surface (or by a projection of a straight connection of said edges in the radial direction on the platform surface) and in the circumferential direction by the suction or pressure side is limited, is referred to in this document as "blade intermediate strip".
  • the width of the blade intermediate strip in the circumferential direction is called the "pitch" of the blade grid. In particular, it can be measured as the distance of the leading edges of respectively adjacent blades in the circumferential direction on the platform surface.
  • the depth of the blade clearance in the axial direction ie the distance of the leading edges of the blades from their trailing edges measured parallel to the intended axis of rotation of the turbomachine, is referred to as "grid width".
  • a fluid flow guided through a flow channel is regularly influenced by the surfaces of the side walls. Flow layers that run close to these surfaces are deflected more strongly because of their lower velocity than laterally from the sidewalls flow layers. This creates a secondary flow, which is superimposed on an axial main flow and which leads in particular to eddies and pressure losses.
  • contouring in the form of elevations and / or depressions is frequently introduced into the side walls.
  • a flow channel with a side wall which has a radial recess in the region of the leading edges of the blades. This extends in the axial direction over the majority of the flow channel and ends only shortly before or even behind the trailing edges. This is intended to locally increase the surface area of the flow area between the inflow and outflow edges, which should improve the efficiency of the rotor.
  • the EP 2 372 088 A2 discloses an integrally machined turbine blade disk having a rim with upstream and downstream edges between which blades and, in the region of the leading edges of the rotor blades, indentations in the rim surface are arranged.
  • the object of the present invention is to provide an alternative technique for a turbomachine with which secondary flows can advantageously be reduced.
  • a vane grate segment for a vane grille of a turbomachine comprises a platform and at least two (adjacent) airfoils which, in accordance with the above, define an airfoil blade with axial grid width by their upstream and downstream edges on the platform surface.
  • the platform has an upstream platform edge that has a contour with a depression. In the axial direction, this depression extends at most by 10% of the grid width in the blade intermediate strip. In particular, the depression can be arranged completely upstream of the leading edges, ie not project into the blade intermediate strip.
  • the depression is thus to be understood as a local expression of a line which extends from the cross-section of one in the two-dimensional one Platform surface lying (area) depression along the platform edge results.
  • a “depression” is to be understood in this document, a local shape in the platform surface in which this extends to the side facing away from the blades.
  • the term (as well as terms such as “lowered” or the like) is therefore based on an orientation or a coordinate system in which the blades extend from the platform surface "up” and a depression accordingly in the opposite direction (to “down” ) leads.
  • the recess is thus arranged completely in the interior of a surface strip of the platform surface whose downstream boundary extends in the axial direction by at most 10% of the grid width downstream of the leading edges; especially against this limitation, the depression is lowered in all its points (one edge of the depression can not be considered as belonging to it).
  • the recess is preferably formed contiguous. According to a specific embodiment, the downstream boundary extends even no further than (at most) 5% of the grid width downstream of the leading edges of the airfoils.
  • a vane grating segment according to the invention may be in one piece or assembled.
  • the platform may be unitary or comprise two or more parts each of which projects from one of the airfoils, or the platform may be formed as at least one separate component disposed between the airfoils.
  • a platform according to the invention is adapted to abut in the circumferential direction on each side of an airfoil and with the blades (none, one or both of which can be integrally formed on the platform / can) together an inventive blade grid segment according to one of the disclosed in this document Embodiments form.
  • a blade lattice according to the invention comprises at least one blade lattice segment according to the invention according to one of the embodiments disclosed in this document.
  • a turbomachine according to the invention comprises one or more vane grates according to the invention.
  • An inventive blade channel leads through an inventive Blade grating segment according to one of the embodiments disclosed in this document, is thus limited by such a blade grid segment as well as a side wall opposite the platform surface (facing the platform surface).
  • the blade channel is bounded in the circumferential direction by the pressure side of one of the blade leaves of the blade grid segment and by the opposite suction side of the (adjacent) other of the blade leaves.
  • An inventive blade lattice segment, a blade lattice according to the invention, a blade channel according to the invention, a platform according to the invention and a turbomachine according to the invention make it possible to improve the secondary flows and thus to reduce losses in the respective hub or housing area. Thus, a high efficiency of the turbomachine can be achieved.
  • the blade lattice segment or the blade lattice or the flow channel or the platform may in particular be part of a low-pressure turbine.
  • the airfoil may be a vane grille or a blade airfoil, so the airfoils may be each guide or blade airfoils.
  • the platform may be configured to limit a blade channel radially inwardly or radially outwardly through the blade grid segment.
  • the upstream-side platform edge is preferably configured to be (at least substantially) adjacent to another (separate) element (eg, the hub or housing or other airfoil) in the turbomachine. It may be adapted to form a portion of a wall of a gap through which cooling fluid is introduced into the annulus of the turbomachine.
  • the upstream-side platform edge (which may include portions of multiple parts of a multi-part platform) is preferably bounded by the (circumferential) positions of the leading edges of the two airfoils; These boundaries may have a physical character (eg, where the platform ends up in peripheral vision in them) or may be defined or defined only abstractly to define the upstream edge of the platform.
  • the upstream side edge of the platform preferably has an extent (or length) in the circumferential direction that is (substantially) equal to the pitch.
  • an embodiment of the present invention has been found in which the recess along the upstream platform edge (preferably continuously) over at least 50% of the pitch extends.
  • the recess has a positive distance (> 0) from the pressure side of one airfoil and / or from the suction side of the other airfoil so that it does not touch the respective side.
  • the recess may be the same or different distances from the two blades.
  • the distance from the recess to the leading edge of one blade e.g., at the pressure side of which the blade clearance area is adjacent
  • the distance between the depression and the leading edge of the other blade may be greater or less than the distance between the depression and the leading edge of the other blade.
  • an embodiment of the present invention has proven to be advantageous in which the inflow-side platform edge (or its contour) is formed asymmetrically to its radial center axis, ie a radial axis which extends through the center of the inflow-side platform edge.
  • the platform surface comprises a surface area disposed between the recess and a pressure side of (a first one of) the airfoils.
  • a surface area is detected by the upstream platform edge.
  • the contour of the inflow-side platform edge then includes an edge of the surface area, which is referred to below as the "pressure-side" surface area.
  • an edge portion of the inflow-side platform edge, in which it detects the pressure-side surface area preferably extend in the circumferential direction over at least 10% or at least 20% of the pitch.
  • Each point of the recess is preferably lowered in comparison with each point of the pressure-side surface area (in the radial direction).
  • the platform surface may include a surface area disposed between the recess and a suction side of the other (second) of the airfoils.
  • a surface area of the upstream platform edge detected.
  • the contour of the inflow-side platform edge then includes an edge of the surface area, which is referred to below as the "suction-side" surface area.
  • an edge portion of the inflow-side platform edge, in which it detects the suction-side surface area preferably extend in the circumferential direction over at least 10% or at least 20% of the pitch.
  • Each point of the recess is preferably lowered in comparison with each point of the suction-side surface area (in the radial direction).
  • the platform surface therefore comprises both a pressure-related and a suction-side surface section (possibly with the other properties mentioned).
  • an edge section in which the inflow-side platform edge detects the suction-side surface area, is larger than an edge section of the inflow-side platform edge, in which it detects the suction-side area area; in another variant, it is reversed, and in another embodiment, both sections are the same size.
  • the platform surface may be formed without contours.
  • FIG. 1 is schematically an exemplary (developed) embodiment of a Shovel segment 1 according to the invention shown in plan view; the viewing direction corresponds to the radial direction (outward or inward, depending on whether the platform 10 is part of an outer or an inner side wall).
  • the airfoil segment comprises adjacent airfoils 20, 30 and a platform 10 according to the invention, which has a platform surface 12, a platform edge 10a (based on the intended axial main flow direction X), and a downstream platform edge 10b.
  • the upstream platform edge 10a may comprise portions of multiple parts of a multi-part platform. It is limited by the (circumferential) positions of the leading edges 23, 33 of the two blades 20, 30; In particular, the extent (or length) of the inflow-side platform edge 10a in the circumferential direction is thus equal to the pitch t.
  • the pressure side 21 of one airfoil 20 and the suction side 32 of the other airfoil 30 delimit in the circumferential direction U of the associated airfoil a blade intermediate strip 11; in the axial direction, this blade intermediate strip is delimited by a straight connection 11a of the leading edges 23, 33 on the one hand and by a corresponding connection 11b of the trailing edges 24, 34 of the blades 20, 30 in the plan view (ie, a corresponding projection) along the platform.
  • the platform surface has a depression 13 which is captured by the upstream platform edge 10a.
  • the upstream platform edge 10a In cross-section (along a plane to which the intended axis of rotation is normal) it follows that the upstream platform edge 10a has a contour with the depression 13 (in the form of a depression) (not directly shown in the figure, but implied by it) ,
  • the depression is arranged completely upstream of the blade intermediate strip 11 in the illustrated example. It extends along the upstream platform edge 10a contiguous (ie, continuous) over an extent d that is greater than 50% of the pitch t.
  • an outflow-side boundary 13b of the recess 13 has a changing with its course axial distance to the inflow-side platform edge 10a;
  • a downstream boundary of the recess 13 could extend substantially without axial deviations in the circumferential direction U, in a radial projection on the Platform surface so parallel to the upstream platform edge 10 a (not shown).
  • the platform surface has a pressure-side surface portion 14 disposed between the pressure side 21 of the airfoil 20 and the recess 13 and approaching the front platform edge 10a, and a suction-side surface portion 15 located between the suction side 32 of the airfoil 30 and the recess 13 and approaching the front platform edge ,
  • the depression 13 is completely lowered in the radial direction in comparison with each point of the pressure-side section 14 and each point of the suction-side section 15 (which is again not visible in the figure as a plan view due to the illustration).
  • the pressure-side surface portion 14 reaching toward the front platform edge 10a extends in a continuous edge portion 14a along the front platform edge 10a.
  • the suction-side portion 15 extending to the front platform edge extends in a continuous edge portion 15a along the front platform edge 10a. I'm in the FIG. 1 As shown, the edge portion 15a is smaller than the edge portion 14a.
  • the inflow-side platform edge 10a is asymmetrical to its (not shown in the figure) radial central axis, ie to a radial axis which passes through the center of the inflow-side platform edge 10a.
  • FIG. 2 schematically shows a developed alternative embodiment of a blade grid segment 1 according to the invention in plan view. It points like that in FIG. 1 Shovel segment 1 shown blade airfoils 20, 30 and a platform 10 according to the invention with a (based on the intended, axial main flow direction X) upstream platform edge 10a and a downstream platform edge 10b on.
  • the platform surface of the platform 10 of the airfoil segment 1 ' comprises a depression 13' running along the upstream side edge 10a of the platform 10a which is on the upstream side.
  • a contour in the inflow-side platform edge 10a results in the depression 13 'in the form of a contiguous depression.
  • a downstream boundary 13'b of the recess 13 ' also has in the FIG. 2 Example shown a changing with its course distance in the axial direction to the upstream platform edge 10a on;
  • a downstream boundary of the recess extends substantially without axial deviations in the circumferential direction (ie, in the projection on the platform surface parallel to the inflow-side platform edge 10a).
  • the recess 13 ' is in the interior of a surface strip of the platform surface 12' is arranged, the downstream boundary extends in the circumferential direction and the axial distance a downstream of the leading edges 23, 33 of the blades 20, 30.
  • a ⁇ 0.1 g where g is the axial grid width.
  • the depression 13 'thus protrudes at most 10% of the axial grid width g into the blade intermediate strip 11.
  • the platform surface has a pressure-side surface portion 14 'disposed between the pressure side 21 of the airfoil 20 and the recess 13' and reaching the front platform edge 11, and a suction side disposed between the suction side 32 of the airfoil 30 and the recess 13 'and approaching the front platform edge Surface portion 15 'on.
  • the recess 13 ' is compared to each point of the pressure-side portion 14' and each point of the suction-side portion 15 'completely lowered in the radial direction (which is again not visible in the figure due to the representation as a plan view).
  • the pressure-side surface portion 14 'extending toward the front platform edge extends continuously along the front platform edge 10a in an edge portion 14'a.
  • the suction-side portion 15 'extending to the front platform edge extends continuously along the front platform edge 10a in an edge portion 15'a.
  • the edge portion 15'a is larger than the edge portion 14'a; in a particular embodiment, the edge portion 15'a may be at least 1.5 times or even at least twice as large as the edge portion 14'a.
  • a blade lattice segment 1, 1 'for a blade lattice of a turbomachine comprising a platform 10 and at least two blades 20, 30, by their arrival and trailing edges 23, 33, 24, 34 on the platform surface, an intermediate blade strip 11 with axial grid width Determine g.
  • An upstream side Platform edge 10a has a contour with a recess 13, 13 '. In the axial direction, this depression 13, 13 'extends at most by 10% of the grating width g in the blade intermediate strips 11.

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  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP17159076.3A 2017-03-03 2017-03-03 Contournage d'une plate-forme de grille d'aube Active EP3369892B1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP17159076.3A EP3369892B1 (fr) 2017-03-03 2017-03-03 Contournage d'une plate-forme de grille d'aube
ES17159076T ES2819128T3 (es) 2017-03-03 2017-03-03 Contorneo de una plataforma de una rejilla de paletas
US15/908,401 US10648339B2 (en) 2017-03-03 2018-02-28 Contouring a blade/vane cascade stage

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP17159076.3A EP3369892B1 (fr) 2017-03-03 2017-03-03 Contournage d'une plate-forme de grille d'aube

Publications (2)

Publication Number Publication Date
EP3369892A1 true EP3369892A1 (fr) 2018-09-05
EP3369892B1 EP3369892B1 (fr) 2020-08-19

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EP17159076.3A Active EP3369892B1 (fr) 2017-03-03 2017-03-03 Contournage d'une plate-forme de grille d'aube

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US (1) US10648339B2 (fr)
EP (1) EP3369892B1 (fr)
ES (1) ES2819128T3 (fr)

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GB202004925D0 (en) * 2020-02-13 2020-05-20 Rolls Royce Plc Aerofoil assembly and method

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WO1998044240A1 (fr) * 1997-04-01 1998-10-08 Siemens Aktiengesellschaft Structure superficielle pour la paroi d'un canal d'ecoulement ou d'une aube de turbine
EP1126132A2 (fr) 2000-02-18 2001-08-22 General Electric Company Paroi radiale profilée pour compresseur
US20080232968A1 (en) * 2006-02-27 2008-09-25 Honeywell International, Inc. Non-axisymmetric end wall contouring for a turbomachine blade row
DE102010007985A1 (de) * 2009-02-27 2010-09-02 Alstom Technology Ltd. Dampfturbine und Verfahren zum Extrahieren von Feuchtigkeit aus einer Dampfturbine
EP2372088A2 (fr) 2009-09-16 2011-10-05 United Technologies Corporation Tranchées de chemin de flux de réacteur à double flux
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EP2487329B1 (fr) 2011-02-08 2013-11-27 MTU Aero Engines GmbH Canal d'aube doté d'une définition de contour de la paroi latérale et turbomachine associé
EP2787172A2 (fr) 2012-08-02 2014-10-08 MTU Aero Engines GmbH Grille d'aubes avec définition de contour de la paroi latérale et turbomachine
EP2696029B1 (fr) 2012-08-09 2015-10-07 MTU Aero Engines AG Grille d'aube avec définition de contour de la paroi latérale et turbomachine
WO2015092204A1 (fr) * 2013-12-18 2015-06-25 Snecma Aube pour roue a aubes de turbomachine et procede de modelisation de celle-ci

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US20180252107A1 (en) 2018-09-06
ES2819128T3 (es) 2021-04-15
US10648339B2 (en) 2020-05-12
EP3369892B1 (fr) 2020-08-19

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