EP1724440B1 - Procédé pour optimiser l'écoulement dans des turbomachines multi-étages - Google Patents

Procédé pour optimiser l'écoulement dans des turbomachines multi-étages Download PDF

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Publication number
EP1724440B1
EP1724440B1 EP05010100A EP05010100A EP1724440B1 EP 1724440 B1 EP1724440 B1 EP 1724440B1 EP 05010100 A EP05010100 A EP 05010100A EP 05010100 A EP05010100 A EP 05010100A EP 1724440 B1 EP1724440 B1 EP 1724440B1
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European Patent Office
Prior art keywords
blade
ring
flow
profiles
blade ring
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Expired - Fee Related
Application number
EP05010100A
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German (de)
English (en)
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EP1724440A1 (fr
Inventor
Andreas Dr. Fiala
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MTU Aero Engines AG
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MTU Aero Engines GmbH
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Application filed by MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Priority to EP05010100A priority Critical patent/EP1724440B1/fr
Priority to ES05010100T priority patent/ES2310307T3/es
Priority to DE502005004938T priority patent/DE502005004938D1/de
Priority to US11/431,365 priority patent/US7758297B2/en
Publication of EP1724440A1 publication Critical patent/EP1724440A1/fr
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Publication of EP1724440B1 publication Critical patent/EP1724440B1/fr
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    • 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
    • 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

Definitions

  • the invention relates to a method for optimizing flow in multi-stage turbomachines, in which the flow of a third of three successive blade rings is optimized, according to the preamble of claim 1.
  • the European Patent EP 0 756 667 B1 protects a generic method of flow optimization in which the relative blade profile positioning between the first and third blade rings is referred to as "clocking".
  • clocking The preferred application here is the vane clocking, ie, the first and third vane rings are vane rings, whereas the second vane ring is a blade ring.
  • the principle of this method is that the flow paths of the trailing edge of the blade profiles of the first blade ring are determined until entering the third blade ring, and the leading edges of the blade profiles of the third blade ring within a predetermined tolerance angle range (25% of the blade pitch angle) relative to the entry positions of Curbs are positioned.
  • Optimal should be a direct / central impact of each caster on the respective leading edge.
  • Each caster descends as a contiguous, turbulent stream from the trailing edge of the blade profile of the first blade ring and, as it passes through the second rotating blade ring, is split into separate pieces moving side-by-side on particular webs.
  • the number of tracks corresponds to the circumference of the flow area divided by the number of blades of the first blade ring.
  • the moving pieces of adjacent wake of the first blade ring move in chronological succession.
  • the trailing pieces are averaged over time, so that, computationally, a coherent trailing occurs, which strikes the third blade ring.
  • Another simplifying assumption of the patented method is that the flow of the wake by the second blade ring is to take place only on a flow surface and is not taken into account that the caster also arranges radially differently.
  • the European Patent EP 1 201 877 B 1 also protects a generic method for flow optimization, which is also illustrated by the example of two relative to each other to be positioned vane rings with a coaxially arranged between them blade ring.
  • a generic method for flow optimization which is also illustrated by the example of two relative to each other to be positioned vane rings with a coaxially arranged between them blade ring.
  • the size of the entropy is called. However, it is also said that there may be other parameters that vary in size without specifying them. In any case, select one of the identified zones and direct it to the leading edges of the blade profiles of the third blade ring. The at least one non-selected zone may pass in the blade airfoil space. It is acknowledged that it may be necessary to evaluate various parameters. It may also be necessary, for. For example, it is necessary first to guide the zone of greater entropy and then the zone of smaller entropy to the entry edges and thereby to determine computationally / experimentally which measure leads to an increase in efficiency. Basically, this patent teaches a trial-and-error principle that forces the skilled person to take several different measures.
  • the object of the present invention is to propose a clear, unambiguous method for flow optimization in multi-stage turbomachinery, which offers a higher probability of success, than the known methods.
  • the tolerance angle is ⁇ 10% of the blade pitch angle of the third blade ring.
  • FIG. 1 shows in the form of a diagram the qualitative course of the current density pu (y) or in the incompressible limit the course of the velocity u (y) of a flowing, frictional medium in the region of a component surface, such as the surface of a flow-around blade profile.
  • the coordinate y is selected at least approximately perpendicular to the flow direction and thus also approximately perpendicular to the component surface flowed around.
  • the coordinate y is preferably defined perpendicular to a local tangent of the surface of the blade profile.
  • the velocity u (y) at the component surface is "zero".
  • the current density pu (y) and the velocity u (y) increase according to the course of a continuous, curved curve up to a value ⁇ e u e (y e ), where y e is the value the speed no longer changes through the viscous boundary layer.
  • the value y e corresponds - at least fairly accurately - to the local boundary layer thickness.
  • the frictional, curved current density or velocity course is replaced by a friction-free course with a constant current density ⁇ e u e or speed u e .
  • the component surface is fictively shifted by the value of the displacement thickness ⁇ * (delta star), ie a blade profile is fictionally thickened accordingly.
  • the frictionless flow model the same mass flow must be obtained as for the actual, frictional flow.
  • ⁇ * ⁇ 0 y e 1 - ⁇ u y p e ⁇ u e y e d y .
  • ⁇ e , u e and y e are the corresponding values at the boundary layer edge.
  • ⁇ * is thus the y value whose horizontal line intersects the current density pu (y) in such a way that two equal areas are included below and above the ⁇ * line between it and the velocity curve pu (y).
  • These two surfaces are in Fig. 1 hatched diagonally opposite. The surfaces are bounded laterally by the vertical y-axis and a vertical line by ⁇ e u e (y e ).
  • the line does not have to be a vertical, but can also be inclined to the y-axis by the change in the outer velocity in the flow outside the boundary layers between the pressure and suction side. See the dashed curves in FIG. 1 , Since ⁇ * changes periodically with periodically arranged blade profiles of the upstream rotating blade ring with the time t, is FIG. 1 as a "snapshot" at a given time in a particular place to look at.
  • the time profile of ⁇ * must be determined over at least one period, in each case for the pressure side DS and the suction side SS of the considered blade profile.
  • FIG. 2 shows qualitatively the determination of the so-called obstruction V of the displacement thickness for the pressure side ⁇ * DS and the displacement thickness for the suction side ⁇ * SS of a blade profile.
  • the courses of the displacement thicknesses are respectively plotted positively above the time axis t. It can be seen that the maxima of ⁇ * DS and ⁇ * SS are different in size (high) and offset in time.
  • the time course of the obstruction V results from additive superimposition of the profiles of the displacement thicknesses of ⁇ * DS and ⁇ * SS . Accordingly, the maximum of the obstruction V max is temporally between the time-offset maxima of the displacement thicknesses.
  • the maximum of the obstruction V max is presently to be determined in the region of the trailing edge of a blade profile and can alternatively be determined from the distribution perpendicular to the trailing edge N of the blade profile in the region downstream of the trailing edge.
  • the local blade profile thickness D is added as a further additive quantity to the displacement thicknesses ⁇ * DS and ⁇ * SS .
  • the term "in the area of the trailing edge of a blade profile" is intended to mean that the location for determining the maximum obstruction near the trailing edge can be selected within the airfoil, directly at the trailing edge, or near the trailing edge downstream of the airfoil. Ultimately, it is important that the further path of the maximum of the obstruction is determined correctly.
  • FIG. 3a shows a longitudinal section through a bladed flow channel with a static housing 46 and with a rotating hub 47.
  • Leitschaufelkränze 48, 50 are arranged with the hub 47 rotate blade rings 49,51.
  • curved contour is shown in each case. Within the flow channel, three more curved lines can be seen. These are the intersections of three rotational flow surfaces ⁇ 2 ⁇ ⁇ . ⁇ 6 ⁇ ⁇ and ⁇ 10 ⁇ ⁇ with the selected, axial-radial cutting plane.
  • the stream surfaces correspond to the spatial trajectories of selected "fluid particles".
  • FIG. 3b shows the implementation of the process principle on the hardware, ie at flow arranged in series blade rings.
  • the flow takes place here from left to right, ie from the blade ring 1 to the blade ring 3.
  • three adjacent blade rings 1 to 3 are considered, of which the first blade ring 1 and the third blade ring 3 belong to the same unit “stator” or "rotor”
  • the second blade ring 2 belongs to the respective complementary unit “rotor” or “stator”.
  • the blade rings 1 and 3 belong to the stator, ie be Leitschaufelkränze.
  • the blade ring 2 should belong to the rotor, ie be a blade ring.
  • the representation according to FIG. 3b For the sake of better clarity, only the blade profiles 18, 19, 20 of the blades 4, 5, 6, 20 are shown on a specific flow surface, ie in a flow area cut.
  • the leading edges of the blade profiles 18,19,20 carry the reference numerals 25,26,27, the exit edges the reference numerals 35,36,37.
  • the upstream blade profiles 18 generate so-called wake N, ie flow areas with turbulence and at reduced speed in the desired flow direction due to friction.
  • each caster N has a circumferential and a meridional component, which in turn can be composed of an axial and a radial component, so that each caster N enters the region of the moving second blade ring 2 and of its successive blade profiles 19 in Separate pieces is divided, which move through the flow channels between the blades 5 and interact with the pressure and suction side boundary layers of the blade profiles 19 in correlation.
  • the blade profiles 19 are in the region of the exit edges 36 there to detect periodically occurring maxima of the obstruction V max locally and temporally in the procedure according to FIG FIGS.
  • the further path of the respective maximum of the obstruction V max is to be traced into the region of the entry edges 27 of the blade profiles 20 of the third, static blade ring 3.
  • the maximum of the obstruction V max should strike an entrance edge 27 within a predetermined tolerance angle ⁇ wt.
  • This tolerance angle is for example ⁇ 15% of the blade pitch angle wt of the third blade ring 3, ie it extends on both sides of the leading edge 27 by 15% in the circumferential direction.
  • the entire angular range is thus 30% of the blade pitch angle wt of the third blade ring 3. If the measurements or calculations show that the maxima of the obstruction V max actually hit the inlet edges of the third blade ring 3 within the predetermined tolerance angle ⁇ wt, the desired flow optimization has been achieved ,
  • the closest measure is likely to be a relative rotation of the blade rings 1 and 3, ie a relative, limited angular movement in the circumferential direction about the longitudinal center axis of the blade rings. After optimization, it must be ensured that the relative positioning during removal and installation or during operation is not unintentionally changeable.
  • Another obvious measure is the axial displacement of at least one of the blade rings 1,2,3, but preferably the blade ring 1 relative to the blade ring 2 to move axially. The same effect is achieved by axial displacement of the blade profiles relative to their carrier, ie relative to the disc, the hub, the shroud, etc. This is usually already associated with further structural changes.
  • the present optimization method will typically not be limited to only one radial flow area, i. in a stream area intersection, but in several, distributed over the radial extent of the airfoil stream surface sections is performed.
  • FIG. 4 shows in axial view on stator trailing edges, the three blades 7,8,9, which emanate from a common foot area, but extend differently over their radial height.
  • the blade 7 shown in solid lines runs straight and radially, ie, rather conventionally “threaded”, ie the profile cuts are placed at the same circumferential position at each trailing edge.
  • the blade 8 shown in dashed lines runs straight, but with an inclination in the circumferential direction. This is also called “lean”.
  • the dot-dashed blade 9 has a curvature in the circumferential direction, a so-called “bow” on. In fact, with such changes, a relative circumferential displacement of the radially superposed profile cuts is achieved.
  • FIG. 5 shows a view of two rotor blades 10,11 in the circumferential direction.
  • the vane 10 shown in solid lines with the leading edge 28 and the trailing edge 38 has a trapezoidal, more conventional outline.
  • the blade 11 shown in dashed lines has an axially curved leading edge and an axially curved exit edge 39 in the same direction. This is also referred to as "axial bow” or “sweep” and causes primarily a relative displacement of the profile sections in the axial direction.
  • FIG. 6 shows a view of two blades 12,13 in the circumferential direction.
  • the blade 12 shown in solid lines with the inlet edge 30 and the outlet edge 40 corresponds in its trapezoidal, conventional outline of the blade 10 FIG. 5
  • the blade 13 shown in dashed lines has the blade root 12 and the blade tip together with the blade 12.
  • Their leading edge 31 and their trailing edge 41 are bent in opposite directions outwards, so that a bulbous blade outline is formed. This measure is also referred to as "barreling".
  • the axial length of the profile sections is primarily increased, wherein the magnification in the region of the average radial height is most pronounced.
  • only another arbitrary profile cut can be common.
  • FIG. 7 shows a profile section through two blades 14,15 with the same blade profiles 21,22 in different position.
  • the entry edges are designated 32.33, the exit edges 42.43.
  • the blade profile 22 shown in dashed lines is intended to be rotated relative to the blade profile 21 shown in solid lines around the threading axis (not reproduced here).
  • the leading edge 32 and the trailing edge 42 of the blade 21 are more circumferentially offset than the leading edge 33 and the trailing edge 43 of the blade 22. This measure is also referred to as "twisting". With the rotation, both the inflow and the outflow of such a blade lattice changes directionally.
  • FIG. 8 finally shows a profile section through two blades 16,17 with the same inflow and different outflow.
  • the two blade profiles 23 and 24 have a common leading edge 34 and a common "nose contour".
  • the blade profile 23 shown in solid lines causes by a stronger profile curvature and a stronger flow deflection up to its trailing edge 44.
  • the blade profile 24 shown in dashed lines deflects the flow less up to its trailing edge. This measure is also known as "vortexing".

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

Claims (5)

  1. Procédé d'optimisation de l'écoulement dans des turbomachines à plusieurs étages, en particulier dans des turbines à gaz, où le flux de la troisième (3) de trois couronnes d'aubes consécutives (1, 2, 3) est optimisé, la première (1) et la troisième (3) couronnes d'aubes ayant le même nombre d'aubes et donc le même pas angulaire d'aubage (wt) et étant toutes les deux disposées coaxialement contre la même unité de rotor ou de stator, la deuxième couronne d'aubes (2) étant disposée coaxialement contre l'autre unité de rotor ou de stator entre la première (1) et la troisième couronne d'aubes (3), de telle manière qu'une rotation relative est présentée en fonctionnement entre la deuxième couronne d'aubes (2) et les deux autres couronnes d'aubes (1, 3), et un état de fonctionnement (point de fonctionnement) étant sélectionné quant aux paramètres de service, avec une part temporelle élevée de la durée de fonctionnement de la turbomachine, des maxima d'obstruction (Vmax) périodiquement formés étant induits pour cet état de fonctionnement au niveau des arêtes de sortie (36) de profils d'aubes (19) de la deuxième couronne d'aubes (2) par des post-écoulements (N) partant de profils d'aubes (18) de la première couronne d'aubes (1) et se déplacent de la zone des arêtes de sortie (36) des profils d'aubes (19) de la deuxième couronne d'aubes (2) vers la zone des arêtes d'entrée (27) de profils d'aubes (20) de la troisième couronne d'aubes (3), étant dirigés vers les arêtes d'entrée (27) des profils d'aubes (20) de la troisième couronne d'aubes (3) à l'intérieur d'un angle de tolérance (Δwt) de ± 15 % du pas angulaire d'aubage (wt) de la troisième couronne d'aubes (3), les positions et/ou les géométries de profils d'aubes (18, 19, 20) d'au moins une des trois couronnes d'aubes (1, 2, 3) étant si besoin modifiées à cet effet, caractérisé en ce que, pour la détermination de grandeur de l'obturation (V) au niveau de l'arête de sortie (36) d'un profil d'aube (19) de la deuxième couronne d'aubes (2), l'épaisseur de déplacement δ* (Deltaster) pour le côté de pression (DS) et le côté d'aspiration (SS) du profil d'aube est calculée suivant la formule δ * = ∫ 0 y e ( 1 - ρu y ρ e ⁢ u e y e ) dy
    Figure imgb0009
    où p représente la densité du fluide qui s'écoule, u la vitesse du fluide qui s'écoule, y une coordonnée perpendiculaire à une ligne de référence au niveau de l'arête de sortie (36) et e un indice pour la limite entre le flux perturbé et non perturbé par la couche limite, et en ce que la grandeur de l'obturation (V) est calculée comme somme des épaisseurs actuelles de déplacement δ*DS et δ*SS côté pression et côté aspiration en option avec ou sans l'épaisseur de profil d'aube (D) au niveau de l'arête de sortie (36) du profil d'aube (19) suivant la formule : V = δ ⁢ * DS + δ ⁢ * SS ou V = D + δ ⁢ * DS + δ ⁢ * SS
    Figure imgb0010
  2. Procédé selon la revendication 1, caractérisé en ce que l'angle de tolérance (Δwt) est de ± 10 % du pas angulaire d'aubage (wt) de la troisième couronne d'aubes (3).
  3. Procédé selon la revendication 1 ou 2, caractérisé en ce qu'il est exécuté de manière répétée pour des hauteurs d'aubes différentes, autrement dit pour plusieurs sections de profil d'aube, du pied au bout de l'aube.
  4. Procédé selon l'une des revendications précédentes, caractérisé en ce qu'il est répété jusqu'à ce que les profils d'aubes de toutes les couronnes d'aubes mobiles et/ou fixes d'un composant de turbomachine, tel qu'un compresseur haute pression ou une turbine basse pression d'une turbine à gaz, présentent un positionnement et/ou un contour optimal pour l'écoulement dans une ou plusieurs sections de profil d'aube.
  5. Procédé selon l'une des revendications précédentes, caractérisé en ce que pour des composants de turbine à gaz, le point dit de ligne aérodynamique (Aerodynamic Design Point - ADP) est appliqué comme état de fonctionnement déterminant, à vitesse nominale (100 %) et avec des conditions de pression totale et de température totale définies.
EP05010100A 2005-05-10 2005-05-10 Procédé pour optimiser l'écoulement dans des turbomachines multi-étages Expired - Fee Related EP1724440B1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP05010100A EP1724440B1 (fr) 2005-05-10 2005-05-10 Procédé pour optimiser l'écoulement dans des turbomachines multi-étages
ES05010100T ES2310307T3 (es) 2005-05-10 2005-05-10 Procedimiento para la optimizacion de la corriente en motores de turbopropulsion de varias fases.
DE502005004938T DE502005004938D1 (de) 2005-05-10 2005-05-10 Verfahren zur Strömungsoptimierung in mehrstufigen Turbomaschinen
US11/431,365 US7758297B2 (en) 2005-05-10 2006-05-10 Method for flow optimization in multi-stage turbine-type machines

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EP05010100A EP1724440B1 (fr) 2005-05-10 2005-05-10 Procédé pour optimiser l'écoulement dans des turbomachines multi-étages

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EP1724440A1 EP1724440A1 (fr) 2006-11-22
EP1724440B1 true EP1724440B1 (fr) 2008-08-06

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ES (1) ES2310307T3 (fr)

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US7967571B2 (en) * 2006-11-30 2011-06-28 General Electric Company Advanced booster rotor blade
FR2925106B1 (fr) 2007-12-14 2010-01-22 Snecma Procede de conception d'une turbine multi-etages de turbomachine
US8540490B2 (en) * 2008-06-20 2013-09-24 General Electric Company Noise reduction in a turbomachine, and a related method thereof
US20090317237A1 (en) * 2008-06-20 2009-12-24 General Electric Company System and method for reduction of unsteady pressures in turbomachinery
US8439626B2 (en) 2008-12-29 2013-05-14 General Electric Company Turbine airfoil clocking
JP5374199B2 (ja) * 2009-03-19 2013-12-25 三菱重工業株式会社 ガスタービン
JP2011241791A (ja) * 2010-05-20 2011-12-01 Kawasaki Heavy Ind Ltd ガスタービンエンジンのタービン
US10287987B2 (en) 2010-07-19 2019-05-14 United Technologies Corporation Noise reducing vane
US9500085B2 (en) * 2012-07-23 2016-11-22 General Electric Company Method for modifying gas turbine performance
US20140068938A1 (en) * 2012-09-10 2014-03-13 General Electric Company Method of clocking a turbine with skewed wakes
US20140072433A1 (en) * 2012-09-10 2014-03-13 General Electric Company Method of clocking a turbine by reshaping the turbine's downstream airfoils
US9435221B2 (en) 2013-08-09 2016-09-06 General Electric Company Turbomachine airfoil positioning
EP2918777B1 (fr) * 2014-03-11 2016-10-26 United Technologies Corporation Méthode d'optimisation d'une aube pour l'atténuation du bruit

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JPS54114618A (en) * 1978-02-28 1979-09-06 Toshiba Corp Moving and stator blades arranging method of turbine
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IT1320722B1 (it) * 2000-10-23 2003-12-10 Fiatavio Spa Metodo per il posizionamento di schiere di stadi di una turbina,particolarmente per motori aeronautici.
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DE10237341A1 (de) * 2002-08-14 2004-02-26 Siemens Ag Modell, Berechnung und Anwendung periodisch erzeugter Kantenwirbel im Turbomaschinenbau

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DE502005004938D1 (de) 2008-09-18
EP1724440A1 (fr) 2006-11-22
ES2310307T3 (es) 2009-01-01
US7758297B2 (en) 2010-07-20

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