EP2665896A1 - Carter intermédiaire - Google Patents

Carter intermédiaire

Info

Publication number
EP2665896A1
EP2665896A1 EP12716196.6A EP12716196A EP2665896A1 EP 2665896 A1 EP2665896 A1 EP 2665896A1 EP 12716196 A EP12716196 A EP 12716196A EP 2665896 A1 EP2665896 A1 EP 2665896A1
Authority
EP
European Patent Office
Prior art keywords
boundary wall
intermediate housing
radially outer
circumferential direction
support ribs
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
EP12716196.6A
Other languages
German (de)
English (en)
Other versions
EP2665896B1 (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
Original Assignee
MTU Aero Engines AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by MTU Aero Engines AG filed Critical MTU Aero Engines AG
Publication of EP2665896A1 publication Critical patent/EP2665896A1/fr
Application granted granted Critical
Publication of EP2665896B1 publication Critical patent/EP2665896B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • 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 an intermediate housing, in particular of turbines of a gas engine, according to the preamble of patent claim 1.
  • 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 Mitteldmckturbine 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.
  • an intermediate housing 14 with a transition flow channel 33 extends to the flow, which leaves the high pressure turbine 11 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 rear edge 17, which is also referred to as a flow outlet edge, and via side walls 18.
  • transition flow channel 33 can (see Fig. 1) upstream of the support ribs 15 in the region of an entry into the transitional flow channel 33 or in the region of a leading edge 34 of the intermediate housing 14 radially outwardly in the same open a cavity 19, through which a small amount of cooling air 21a, which deals with the high pressure
  • CONFIRMATION COPY 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 21 c. 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 area of entry into the cavity 19 is below the pressure of the cooling air 21b in the secondary air area 21d of the annulus.
  • FIG. 2 shows a dimensionless circumferential direction u / t, where t corresponds to the support rib pitch in the circumferential direction u.
  • the pressure fields of the pressure increase ⁇ on the circumferential position of the support ribs 1 and the pressure drop ⁇ at the circumferential position between adjacent support ribs 15 respectively upstream of the leading edges 16 of the support ribs 15 are shown in dashed lines in the cavity 19 so that Further, the pressure fluctuation of FIG. 2 in the cavity leads to a higher pressure gradient between the gas flow 20 and the cooling air flow 21b, which ultimately increases the leakage and degraded efficiency the turbomachine leads.
  • 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.
  • Prior art known turbomachine in the region of an intermediate housing and thus flow channel between two turbine components; 2 shows a detail of the arrangement of Figure 1 in the radial direction.
  • FIG. 3 is a highly schematic, fragmentary longitudinal section through a flow machine in the region of an intermediate housing according to the invention, which is positioned between two turbine components;
  • Fig. 4 is a diagram for illustrating the invention.
  • Fig. 5 is another diagram for illustrating the invention.
  • 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 skilled in the art and has already been described in connection with FIG.
  • 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 sselluftströmungska- channel 19 cooling air flow can be improved in the guided by the transitional flow channel 33 of the intermediate housing 14 gas flow, namely in an inlet region of the transitional flow channel 33 upstream of in 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 a turbine component positioned upstream of the transitional flow channel 33, designed as a high-pressure turbine 11 in the exemplary embodiment shown.
  • FIG. 3 shows the cooling air flow channel 19 from the radially outward direction the transitional flow channel 33 opens, namely upstream of support ribs 15 which are positioned in the transitional flow channel 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 can be located at least in a section upstream of the support ribs 15 be provided with a circumferentially changing contour.
  • the radially outer boundary wall 24 of the transitional flow channel 33 at least in a transition section between the front edge 34 of the intermediate housing 14 and the transitional 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 can also extend into a region downstream of the leading edges 16 of the support ribs 15, FIG. 3 showing two contours formed at different circumferential positions u / t 24 and 24 'for the radially outer boundary wall of the transitional flow channel 33 shows.
  • 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 section 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 minimal Radius of curvature in the circumferential direction u or ut 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 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 from axial positions, the axial position of the boundary wall point 26 changes continuously or continuously 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.
  • FIG. 3 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 1, whereas the contour shown in Fig. 3 24 'of the same contour approximately in a circumferential position Division between two adjacent support ribs 15 corresponds. Further details regarding the offset of the axial position and the radial position of the boundary wall point 26 with a minimum radius of curvature in the circumferential direction u or u / t are described below with reference to FIG. 4. In Fig. 4 is on the horizontal axis an absolute value ⁇ / X K S between the axial distance ⁇ (see Fig.
  • the region 28 of FIG. 4 visualizes a preferred range of validity for the ratio ⁇ / XKS and / or Ar / x ⁇ s changing in the circumferential direction u or ut and thus the offset of the axial position u and u / t changing in the circumferential direction u or u / t / or the radial position of the boundary wall 26 with a minimum radius of curvature.
  • the ratios Ax / x K s and ⁇ / XS are up to 40%.
  • the ratio changes in the circumferential direction u or u / t
  • the curve 29 within the range 28 visualizes the preferred circumferentially varying ratio AX / XS and thus the circumferentially varying offset of the axial position of the minimum crimp radius limiting wall point 26, where the offset of the axial position is in the range of half Division between two adjacent support ribs is the largest and the ratio ⁇ / ⁇ ⁇ 5 is about 20%.
  • the curve 30 within region 28 illustrates the preferred circumferentially varying ratio Ar / x K s and thus the circumferentially varying offset of the radial position of the boundary wall 26 with minimum radius of curvature, with approximately half pitch between adjacent support ribs
  • Ratio ⁇ KS is about 2.5% and the offset of the radial position in the region 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 Ax / x K s and Ar / x K s respectively change continuously and continuously preferably not linear.
  • FIG. 5 visualizes the effect of contouring the radially outer boundary wall 24 of the transition flow channel 33, FIG. 5 showing on the horizontally extending axis a ratio (pp m ) / pm between the difference (pp m ) of the static pressure p of FIG Gas flow is plotted in the transitional flow channel 14 and the mean value p m of this static pressure and the mean value p m , and wherein on the vertical axis, the dimensionless circumferential direction u / t is plotted.
  • the curve 31 of FIG. 5 corresponds to a self-adjusting according to the prior art, the course of the ratio (pp m) / p m, and the curve 32 the self-adjusting according to the invention the course of the ratio (pp m) / p m. 5, it can be seen that with the invention, an improved, uniform pressure profile of the static pressure in the circumferential direction can be provided, whereby the formation of a secondary flow in the mouth portion of thede Kunststoffströ- mung 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 3 can be ensured, whereby the efficiency of the flow machine can be improved. Furthermore, the flow in the transitional flow passage 33 between adjacent support ribs 15 can be improved.

Landscapes

  • 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)

Abstract

L'invention concerne un carter intermédiaire (4), notamment pour turbines (11, 13) d'un mécanisme d'entraînement à gaz, comprenant une paroi de délimitation (23) située à l'intérieur dans le sens radial et une paroi de délimitation (24, 24') située à l'extérieur dans le sens radial, un canal d'écoulement de transition (33) qui est formé par les parois de délimitation (23, 24, 24') et dans lequel est positionnée au moins une nervure-support (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), qui guident un flux gazeux parcourant le canal d'écoulement de transition (33). La paroi de délimitation (24) située à l'extérieur dans le sens radial comporte, au moins dans un segment situé en amont de la nervure-support (15), un contour variant dans le sens périphérique.
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 Not-in-force 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 true EP2665896A1 (fr) 2013-11-27
EP2665896B1 EP2665896B1 (fr) 2015-06-10

Family

ID=45999502

Family Applications (1)

Application Number Title Priority Date Filing Date
EP12716196.6A Not-in-force 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
EP2990600B1 (fr) 2014-08-29 2017-06-07 MTU Aero Engines GmbH Ensemble de turbine à gaz
DE102017222193A1 (de) 2017-12-07 2019-06-13 MTU Aero Engines AG Turbomaschinenströmungskanal

Family Cites Families (13)

* Cited by examiner, † Cited by third party
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
US7690890B2 (en) * 2004-09-24 2010-04-06 Ishikawajima-Harima Heavy Industries Co. Ltd. Wall configuration of axial-flow machine, and gas turbine engine
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
ATE497087T1 (de) 2007-08-06 2011-02-15 Alstom Technology Ltd Spaltkühlung zwischen brennkammerwand und turbinenwand einer gasturbinenanlage
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

Non-Patent Citations (1)

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Title
See references of WO2012097798A1 *

Also Published As

Publication number Publication date
US20130064657A1 (en) 2013-03-14
US9382806B2 (en) 2016-07-05
EP2665896B1 (fr) 2015-06-10
DE102011008812A1 (de) 2012-07-19
WO2012097798A1 (fr) 2012-07-26

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