EP2665896B1 - Carter intermédiaire d'une turbine à gaz comprenant une paroi de limite extérieure en amont d'une entretoise avec un contour variable en direction circonférentielle pour minimiser les pertes de flux secondaires - Google Patents

Carter intermédiaire d'une turbine à gaz comprenant une paroi de limite extérieure en amont d'une entretoise avec un contour variable en direction circonférentielle pour minimiser les pertes de flux secondaires Download PDF

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Publication number
EP2665896B1
EP2665896B1 EP12716196.6A EP12716196A EP2665896B1 EP 2665896 B1 EP2665896 B1 EP 2665896B1 EP 12716196 A EP12716196 A EP 12716196A EP 2665896 B1 EP2665896 B1 EP 2665896B1
Authority
EP
European Patent Office
Prior art keywords
boundary
wall
housing
boundary wall
radially outer
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.)
Active
Application number
EP12716196.6A
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German (de)
English (en)
Other versions
EP2665896A1 (fr
Inventor
Martin Hoeger
Inga Mahle
Jochen Gier
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
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Publication of EP2665896A1 publication Critical patent/EP2665896A1/fr
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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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/14—Casings modified therefor
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • 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
    • 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
    • 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
    • F05D2270/17—Purpose of the control system to control boundary layer

Definitions

  • the invention relates to a housing of a gas engine according to the preamble of patent claim 1, as this from the US 2008/0276621 A1 and WO 2009/019282 A2 is known.
  • a multi-shaft turbomachine such as a multi-shaft gas engine has a plurality of compressor components, at least one combustion chamber and a plurality of turbine components.
  • a two-shaft gas engine has a low-pressure compressor, a high-pressure compressor, at least one combustion chamber, a high-pressure turbine and a low-pressure turbine.
  • a three-shaft gas engine has a low-pressure compressor, a medium-pressure compressor, a high-pressure compressor, at least one combustion chamber, a high-pressure turbine, a medium-pressure turbine and a low-pressure turbine.
  • Fig. 1 shows a highly schematic section of a multi-shaft gas engine in the region of a rotor 10 of a high-pressure turbine 11 and a rotor 12 of a low-pressure turbine 13. Between the high-pressure turbine 11 and the low-pressure turbine 13 extends between an intermediate housing 14 with a transitional flow channel 33 to the flow, which the high-pressure turbine 11 leaves to supply the low-pressure turbine 13, wherein in the transitional flow channel 33 at least one support rib 15 is positioned.
  • the support rib 15 is a stator-side component, which carries the flow flowing through the transition flow channel 33.
  • a flow-guiding support rib 15 has a front edge 16, which is also referred to as a flow inlet edge, via a trailing edge 17, which is also referred to as a flow outlet edge, and via side walls 18.
  • transitional flow channel 33 can (see Fig. 1 ) open upstream of the support ribs 15 in the region of an entry into the transitional flow channel 33 and in the region of a front edge 34 of the intermediate housing 14 radially outwardly in the same a cavity 19 through which a small amount of cooling air 21 a can emerge, which deals with the high-pressure turbine 11 leaving gas flow 20 mixed.
  • This cavity 19 is located between the NDT housing and the intermediate housing 14, which is sealed with a seal 21c. Only a weak leakage flow 21b flows through this seal 21c, since the NDT housing and the intermediate housing 14 can not be firmly connected to each other.
  • the static pressure of the gas flow 20 in the region of entry into the cavity 19 is below the pressure of the cooling air 21b in the secondary air region 21d outside the annulus.
  • FIG. 2 can be removed, arises in the from the prior art according to Fig. 1 known turbomachine upstream of the leading edges 16 of the support ribs 15 due to a blocking of the gas flow flowing through the transitional flow channel 33 at circumferential positions on which the support ribs are positioned, a pressure increase + .DELTA.p of the static pressure, whereas according Fig. 2 on circumferential positions between adjacent support ribs 15 sets a pressure drop - ⁇ p of the static pressure.
  • Fig. 2 is shown a dimensionless circumferential direction u / t, where t corresponds to the support rib pitch in the circumferential direction u.
  • the present invention is based on the problem to provide an intermediate housing, by means of which the efficiency can be increased.
  • the radially outer boundary wall has a contour which changes in the circumferential direction at least in a section upstream of the support rib.
  • the present invention relates to the field of multi-shaft turbomachinery, in particular multi-shaft gas engines, with several compressor components and several turbine components.
  • the basic structure of such a turbomachine is familiar to the person mentioned here and has already been in connection with Fig. 1 described.
  • the present invention now relates to details of an intermediate housing 14 of such a turbomachine, by means of which the entry of a guided in a cooling air flow passage 19 cooling air flow can be improved in the outflow passage 33 of the intermediate housing 14 guided gas flow, namely in an inlet region of the transitional flow channel 33 upstream of the transitional flow channel 33 positioned support ribs 15th
  • the invention is applicable both to an intermediate housing 14 of a twin-shaft turbomachine which extends between a high-pressure turbine 11 and a low-pressure turbine 13, and to an intermediate housing of a three-shaft turbomachine which extends between a high-pressure turbine and a medium-pressure turbine or between a medium-pressure turbine and a low-pressure turbine, used.
  • Fig. 3 shows a section of a turbomachine in the region of an intermediate housing 14, a transitional flow channel 33 of this intermediate housing 14 and an upstream of the transitional flow channel 33 positioned, formed in the illustrated embodiment as a high-pressure turbine 11 turbine component, according to Fig. 3 the cooling air flow passage 19 opens radially outward into the transition flow passage 33, namely upstream of support ribs 15 positioned in the transition flow passage 33.
  • the cooling air flow channel 19 is thereby limited in sections by the front edge 34 of the intermediate housing 14.
  • the transitional flow channel 33 is bounded radially inwardly by a stator-side boundary wall 23 and also radially on the outside by a stator-side boundary wall 24.
  • a boundary wall 25 of the high-pressure turbine 11 adjoins the rotor 10 of the high-pressure turbine 11 radially on the outside.
  • the radially outer boundary wall 24 of the transition flow channel 33 at least in a transition section between the front edge 34 of the intermediate housing 14 and the transition flow channel 33 has a circumferentially changing contour.
  • This circumferentially changing contour of the radially outer boundary wall 24 of the transition flow channel 33 may according to Fig. 3 extend to a region downstream of the leading edges 16 of the support ribs 15, wherein Fig. 3 two contours 24 and 24 'formed at different circumferential positions u / t for the radially outer boundary wall of the transitional flow channel 33.
  • the radially outer boundary wall 24 of the transition flow channel 33 has in the inlet region of the transition flow channel 33 upstream of the leading edges 16 of the support ribs 15 via a boundary wall portion or boundary wall 26 with a minimum radius of curvature and thus maximum curvature.
  • the contour of the radially outer boundary wall 24 of the transition flow channel 33 changes in the circumferential direction u or u / t such that an axial position (axial direction x) and / or a radial position (radial direction r) of the boundary wall section or boundary wall point 26 with a minimum radius of curvature in the circumferential direction u or u / t changed.
  • both the axial position and the radial position of the boundary wall point 26 change with a minimum radius of curvature.
  • the axial position of the boundary wall 26 with minimal radius of curvature changes in the circumferential direction u or u / t such that approximately at the circumferential position of the leading edges 16 of the support ribs 15 of this boundary wall 26 in the axial direction x maximum upstream and approximately in a circumferential position half pitch between two adjacent supporting ribs in the axial direction x offset or positioned downstream of maximum. Between these maximum upstream and downstream axial positions, the axial position of the boundary wall 26 gradually changes in the circumferential direction.
  • the radial position of the boundary wall 26 with minimal radius of curvature changes in the circumferential direction u or u / t such that approximately at the circumferential position of the leading edges 16 of the support ribs 15 of this boundary wall 26 in the radial direction r maximum radially outward and approximately in a circumferential position half pitch between two adjacent support ribs 15 in the radial direction r is offset or positioned maximally radially inward. Between these maximum radially inner and radially outer radial positions, the radial position of the boundary wall point 26 changes continuously or continuously in the circumferential direction.
  • contour 24 of the radially outer boundary wall of the transition flow channel 33 corresponds to the contour thereof approximately at the circumferential position of a front edge 16 of a support rib 15, whereas the in Fig. 3 shown contour 24 'of the same contour approximately in a circumferential position half pitch between two adjacent support ribs 15 corresponds.
  • Fig. 4 is on the horizontally extending axis an absolute value ratio ⁇ x / x KS between the axial distance ⁇ x (see Fig. 3 ) Downstream of the axial position and the axial position of the maximum of the upstream boundary wall point 26 with a minimum radius of curvature and the axial distance ⁇ KS (see Fig. 3 ) of a downstream end 27 of the radially outer boundary wall 25 of the upstream of the transition channel 33 positioned turbine component 11 and the front edge 16 of the support ribs 15 applied. Furthermore, in Fig. 4 on the horizontally extending axis an absolute value ratio ⁇ r / x KS between the radial distance ⁇ r (see Fig.
  • x corresponds to KS (see Fig. 3 ) the distance between the downstream end 27 of the radially outer boundary wall 25 of the high-pressure turbine 11 and the front edge 16 of the support ribs 15th
  • the area 28 of the Fig. 4 visualizes a preferred scope for extending u and u / t changing in the circumferential direction ratio Ax / x KS and / or AR / x KS and thus the to u and u / t changing in the circumferential direction of offset of the axial position and / or the radial position of the Boundary wall point 26 with a minimum radius of curvature.
  • the ratios ⁇ x / x KS and ⁇ r / x KS are up to 40%.
  • Curve 29 within region 28 visualizes the preferred circumferentially varying ratio ⁇ x / x KS, and hence the circumferentially varying offset of the axial position of the minimum wall radius limiting wall 26, where, according to curve 29, the axial position offset is in the half pitch range between two adjacent support ribs is largest and the ratio ⁇ x / x KS is about 20%.
  • Curve 30 within region 28 illustrates the preferred circumferentially varying ratio ⁇ r / x KS, and thus the circumferentially varying offset of the radial position of boundary wall 26 with minimum radius of curvature, and at approximately half pitch between adjacent support ribs, the ratio ⁇ r / x KS is about 2.5% and the offset of the radial position in the area of half pitch between two adjacent support ribs is the largest.
  • the offset of the axial position of the boundary wall 26 having the minimum radius of curvature and the offset of the radial position of the boundary wall 26 having the minimum radius of curvature and the above ratios ⁇ x / x KS and ⁇ r / x KS are respectively continuous and continuous, and preferably nonlinear.
  • Fig. 5 visualizes the effect of the contouring according to the invention of the radially outer boundary wall 24 of the transitional flow channel 33 wherein Fig. 5 is plotted on the horizontal axis a ratio (pp m ) / p m between the difference (pp m ) of the static pressure p of the gas flow in the transitional flow channel 14 and the mean value p m of this static pressure and the mean value p m ; vertically extending axis the dimensionless circumferential direction u / t is plotted.
  • the curve 31 of the Fig. 5 corresponds to a state of the art adjusting course of the ratio (pp m ) / p m and the curve 32 according to the invention adjusting the course of the ratio (pp m ) / p m .
  • Fig. 5 It can be seen that with the invention, an improved, uniform pressure distribution of the static pressure in the circumferential direction can be provided, whereby the formation of a secondary flow in the mouth portion of the cooling air flow channel 19 in the transitional flow channel 33 can be effectively counteracted. Thereby, an unhindered entry of the cooling air flow into the transitional flow channel 33 can be ensured, whereby the efficiency of the turbomachine can be improved. Furthermore, the flow in the transitional flow passage 33 between adjacent support ribs 15 can be improved.

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

Claims (6)

  1. Carter (14) d'une turbine à gaz comprenant une paroi de délimitation située radialement à l'intérieur (23) et une paroi de délimitation située radialement à l'extérieur (24, 24'), un canal d'écoulement transitoire (33) formé par les parois de délimitation (23, 24, 24') et dans lequel est positionnée au moins une nervure de soutien (15) qui présente une arête avant (16), une arête arrière (17) ainsi que des parois latérales (18) s'étendant entre l'arête avant (16) et l'arête arrière (17) et conduisant le courant de gaz s'écoulant à travers le canal d'écoulement transitoire (33), la paroi de délimitation située radialement à l'extérieur (24) présentant un contour qui se modifie dans le sens périphérique du moins dans un tronçon en amont de la nervure de soutien (15), et le contour de la paroi de délimitation située radialement à l'extérieur (24) se modifiant de telle façon qu'une position axiale et/ou radiale d'un tronçon de paroi de délimitation ou d'un point (26) de la paroi de délimitation d'un rayon de courbure minimal se modifie dans le sens périphérique,
    caractérisé en ce que
    le carter (14) est un carter intermédiaire (14) de turbines (11,13) de la turbine à gaz,
    - la position axiale du tronçon de paroi de délimitation ou du point (26) de la paroi de délimitation d'un rayon de courbure minimal se modifie dans le sens périphérique de telle façon que, sur la position périphérique d'arêtes avant (16) des nervures de soutien (15), ce point (26) est positionné le plus en amont et approximativement sur une position périphérique d'une demi-graduation entre deux nervures de soutien voisines les plus en aval ; et/ou
    - la position radiale du tronçon de paroi de délimitation ou le point (26) de la paroi de délimitation d'un rayon de courbure minimal se modifie dans le sens périphérique de telle façon qu'approximativement sur la position périphérique d'arêtes avant (16) des nervures de soutien (15), ce point (26) est positionné au maximum radialement à l'extérieur et approximativement sur une position périphérique de demi-graduation entre deux nervures de soutien voisines au maximum radialement à l'intérieur.
  2. Carter (14) selon la revendication 1, caractérisé en ce que la paroi de délimitation (24), située radialement à l'extérieur, du canal d'écoulement transitoire (33) présente, du moins dans un tronçon transitoire entre une arête avant (34) du carter intermédiaire (14) et le canal d'écoulement transitoire (33), un contour qui se modifie dans le sens périphérique.
  3. Carter (14) selon l'une quelconque des revendications précédentes, caractérisé en ce que la valeur du rapport entre la distance axiale de la position axiale aval et de la position axiale la plus en amont du point (26) de la paroi de délimitation d'un rayon de courbure minimal et la distance axiale d'une extrémité aval (27) d'une paroi de délimitation (25) située radialement à l'extérieur d'un composant de turbine (11) positionné en amont du canal d'écoulement transitoire (33) et de l'arête avant (16) des nervures de soutien (15) peut aller jusqu'à 40 %.
  4. Carter (14) selon la revendication 3, caractérisé en ce que ce rapport peut aller jusqu'à 25 %.
  5. Carter (14) selon l'une quelconque des revendications précédentes, caractérisé en ce que la valeur du rapport entre la distance radiale de la position radialement la plus à l'extérieur et de la position radiale radialement à l'intérieur du point (26) de la paroi de délimitation d'un rayon de courbure minimal et la distance axiale d'une extrémité aval (27) d'un paroi de carter (25) située radialement à l'extérieur d'un composant de turbine (11) positionné en amont du canal d'écoulement transitoire (33) et de l'arête avant (16) des nervures de soutien (15) peut aller jusqu'à 40 %.
  6. Carter (14) selon la revendication 5, caractérisé en ce que ce rapport peut aller jusqu'à 5 %.
EP12716196.6A 2011-01-19 2012-01-16 Carter intermédiaire d'une turbine à gaz comprenant une paroi de limite extérieure en amont d'une entretoise avec un contour variable en direction circonférentielle pour minimiser les pertes de flux secondaires Active EP2665896B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102011008812A DE102011008812A1 (de) 2011-01-19 2011-01-19 Zwischengehäuse
PCT/DE2012/000032 WO2012097798A1 (fr) 2011-01-19 2012-01-16 Carter intermédiaire

Publications (2)

Publication Number Publication Date
EP2665896A1 EP2665896A1 (fr) 2013-11-27
EP2665896B1 true EP2665896B1 (fr) 2015-06-10

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EP12716196.6A Active EP2665896B1 (fr) 2011-01-19 2012-01-16 Carter intermédiaire d'une turbine à gaz comprenant une paroi de limite extérieure en amont d'une entretoise avec un contour variable en direction circonférentielle pour minimiser les pertes de flux secondaires

Country Status (4)

Country Link
US (1) US9382806B2 (fr)
EP (1) EP2665896B1 (fr)
DE (1) DE102011008812A1 (fr)
WO (1) WO2012097798A1 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9598981B2 (en) * 2013-11-22 2017-03-21 Siemens Energy, Inc. Industrial gas turbine exhaust system diffuser inlet lip
ES2632613T3 (es) 2014-08-29 2017-09-14 MTU Aero Engines AG Grupo constructivo de turbina de gas
DE102017222193A1 (de) 2017-12-07 2019-06-13 MTU Aero Engines AG Turbomaschinenströmungskanal

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009019282A2 (fr) * 2007-08-06 2009-02-12 Alstom Technology Ltd Installation de turbine à gaz

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Publication number Priority date Publication date Assignee Title
DE19650656C1 (de) 1996-12-06 1998-06-10 Mtu Muenchen Gmbh Turbomaschine mit transsonischer Verdichterstufe
EP1515000B1 (fr) * 2003-09-09 2016-03-09 Alstom Technology Ltd Aubage d'une turbomachine avec un carenage contouré
WO2006033407A1 (fr) * 2004-09-24 2006-03-30 Ishikawajima-Harima Heavy Industries Co., Ltd. Forme de paroi de machine a flux axial et turbomoteur a gaz
US7179049B2 (en) 2004-12-10 2007-02-20 Pratt & Whitney Canada Corp. Gas turbine gas path contour
US8511978B2 (en) 2006-05-02 2013-08-20 United Technologies Corporation Airfoil array with an endwall depression and components of the array
US7594405B2 (en) * 2006-07-27 2009-09-29 United Technologies Corporation Catenary mid-turbine frame design
JP5283855B2 (ja) * 2007-03-29 2013-09-04 株式会社Ｉｈｉ ターボ機械の壁、及びターボ機械
DE102008021053A1 (de) 2008-04-26 2009-10-29 Mtu Aero Engines Gmbh Nachgeformter Strömungspfad einer Axialströmungsmaschine zur Verringerung von Sekundärströmung
DE102008031789A1 (de) * 2008-07-04 2010-01-07 Deutsches Zentrum für Luft- und Raumfahrt e.V. Verfahren und Vorrichtung zur Beeinflussung von Sekundärströmungen bei einer Turbomaschine
DE102008060847B4 (de) * 2008-12-06 2020-03-19 MTU Aero Engines AG Strömungsmaschine
EP2248996B1 (fr) * 2009-05-04 2014-01-01 Alstom Technology Ltd Turbine à gaz
EP2261462A1 (fr) * 2009-06-02 2010-12-15 Alstom Technology Ltd Paroi d'extrémité pour un étage de turbine

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009019282A2 (fr) * 2007-08-06 2009-02-12 Alstom Technology Ltd Installation de turbine à gaz

Also Published As

Publication number Publication date
WO2012097798A1 (fr) 2012-07-26
DE102011008812A1 (de) 2012-07-19
EP2665896A1 (fr) 2013-11-27
US9382806B2 (en) 2016-07-05
US20130064657A1 (en) 2013-03-14

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