EP2179143A2 - Installation de turbine à gaz - Google Patents

Installation de turbine à gaz

Info

Publication number
EP2179143A2
EP2179143A2 EP08786927A EP08786927A EP2179143A2 EP 2179143 A2 EP2179143 A2 EP 2179143A2 EP 08786927 A EP08786927 A EP 08786927A EP 08786927 A EP08786927 A EP 08786927A EP 2179143 A2 EP2179143 A2 EP 2179143A2
Authority
EP
European Patent Office
Prior art keywords
wall
gap
radially
turbine
positive
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
EP08786927A
Other languages
German (de)
English (en)
Other versions
EP2179143B1 (fr
Inventor
Ulrich Steiger
Willy Heinz Hofmann
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.)
General Electric Technology GmbH
Original Assignee
Alstom Technology 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 Alstom Technology AG filed Critical Alstom Technology AG
Publication of EP2179143A2 publication Critical patent/EP2179143A2/fr
Application granted granted Critical
Publication of EP2179143B1 publication Critical patent/EP2179143B1/fr
Active 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
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor

Definitions

  • the present invention relates to a gas turbine plant, in particular for a power plant.
  • Such a gas turbine plant usually comprises a combustion chamber which radially delimits a combustion gas path at least in an annular outlet region with a chamber inner wall and with a chamber outer wall. Furthermore, such a gas turbine plant usually comprises a turbine which radially delimits a turbine gas path at least in a stationary annular inlet area with a turbine inner wall and with a turbine outer wall. In order to avoid excessive component voltages during transient operating conditions of the gas turbine plant, the gas turbine plant can also be provided with a radially inner and / or radially outer axially between the
  • Combustion chamber and the turbine running gap be provided, at which the inner and / or outer chamber wall and the inner and / or outer turbine wall end.
  • a cooling gas supply be provided, which introduces a cooling gas into the turbine gas path and / or in the combustion gas path via the gap.
  • a gas turbine is known from US 2006/0034689 A1, in which a row of guide vanes has on its inflow side a multiplicity of flow guide elements projecting radially into the gas path. With the aid of these flow guide elements, a leakage flow flowing around the blade tips upstream of the rotor blades can be reduced or deflected in the axial direction so as to increase the efficiency of the gas turbine.
  • the invention as characterized in the claims, deals with the problem of providing for a gas turbine plant of the type mentioned in an improved embodiment, which is characterized in particular by an increased efficiency. This problem is solved according to the invention by the subject matter of the independent claim. Advantageous embodiments are the subject of the dependent claims.
  • the invention is based on the general idea of achieving a pressure distribution in the gap influencing the flow of cooling gas by a specific positioning of circumferentially alternating positive and negative steps along the gap.
  • This pressure distribution can be specifically set up so that areas with increased cooling demand are subjected to a higher cooling gas flow than areas with a smaller cooling demand. As a result, the total amount of cooling gas required can be reduced, which ultimately increases the efficiency of the gas turbine plant.
  • the combustion chamber wall or the turbine wall is correspondingly contoured in the end section adjoining the gap.
  • the circumferentially varying contours of the respective chamber wall and the respective turbine wall in the region of the radially inner gap and / or in the region of the radially outer gap can be selected and designed on the basis of different criteria. For example, in order to design the alternately arranged in the circumferential direction arrangement of positive and negative steps at the gap in the circumferential direction distribution of alternating pressure sides and suction sides of
  • Guide vanes are used in a stationary inlet area arranged Leitschaufelsch.
  • This guide blade row arranged in the entry region is usually the so-called “first row of guide blades.”
  • the pressure sides and suction sides of the guide blades bring about this Also upstream to the gap varying pressures in the circumferential direction, which may affect the flow of cooling gas in the gap.
  • these adverse influences can be reduced.
  • a pressure distribution which varies in the circumferential direction during operation of the gas turbine plant at or in the exit region of the combustion chamber can also be taken into account. It has been found that in the exit region of the combustion chamber in the circumferential direction different
  • the pressure distribution or velocity distribution which arises during steady-state operation of the gas turbine can be stationary and could, for example, be due to the burner arranged distributed in the circumferential direction in a multi-burner combustion chamber.
  • the pressures varying in the circumferential direction in the exit region of the combustion chamber also influence the flow of cooling gas through the gap.
  • the disadvantageous influence on the cooling gas flow can be reduced.
  • FIG. 1 shows a greatly simplified axial section of a gas turbine plant in the region of a gap between a combustion chamber and a turbine.
  • a gas turbine plant 1 which is preferably used in a power plant, ie stationary, a combustion chamber 2 and a turbine 3, between which an axial gap 4 is arranged. Furthermore, a cooling gas supply 5 indicated by arrows is provided.
  • a longitudinal center axis or axis of rotation X As a reference for the radial and axial orientation is entered in Fig. 1, a longitudinal center axis or axis of rotation X.
  • the combustion chamber 2 has, at least in an annular outlet region 6, a chamber inner wall 7 and a chamber outer wall 8, which together radially bound a combustion chamber gas path 9 indicated by an arrow.
  • the turbine 3 has, at least in a stationary, ie stator, annular inlet region 10, a turbine inner wall 1 1 and a turbine outer wall 12, which together define a radially indicated by an arrow turbine gas path 13.
  • the turbine 3 in the stationary inlet region 10 can usually have a row of guide vanes 14 with a plurality of guide vanes 15 which are adjacent in the circumferential direction. Since this vane row 14 is the first row of blades impinged by the hot gases of the combustion chamber 2, it is usually also referred to as the first row of stator blades 14.
  • the gap 4 consists in the example shown of a radially inner gap 16 and a radially outer gap 17.
  • the radially inner gap 16 is referred to below as the inner gap 16 or inner gap 16.
  • the radially outer gap 17 is also referred to as the outer gap 17 or outer gap 17.
  • the chamber inner wall 7 and the turbine inner wall 1 1 each end axially.
  • the chamber outer wall 8 and the turbine outer wall 12 each end axially.
  • the cooling gas feed 5 is designed such that it introduces a cooling gas into the turbine gas path 13 or into the combustion gas path 9 via the gap 4 or via the respective sub-gap 16 or 17.
  • the introduction of cooling gas into the gap 4 serves to avoid the entry of hot gases from the combustion gas path 9 or from the turbine gas path 13 through the gap 4 in the areas behind the respective chamber walls 7, 8 and turbine walls 11, 12th
  • the turbine walls 7, 8 may be formed, for example, by heat shield elements or by so-called liners 18.
  • the turbine walls 11, 12 can by platforms 19 and 20, which are formed on the respective blade root radially outside and inside, be formed.
  • alternating positive and negative steps are formed in the axial direction. If we achieve this by means of a corresponding shaping of an end section of the respective turbine wall 11, 12 or of the respective chamber wall 7, 8 adjoining the respective gap 4 or 16 or 17.
  • An end section of the chamber inner wall 7 is designated 21
  • an end section of the chamber outer wall 8 is designated 22
  • an end portion of the turbine inner wall 1 1 is denoted by 23
  • an end portion of the turbine outer wall 12 is denoted by 24.
  • Regions of the end portions 21 to 24 lying in the sectional plane are shown by solid lines, while offset portions of the end portions 21 to 24 thereof are shown by broken lines.
  • the named stages are denoted by the lowercase letters a to d.
  • a denotes a positive step formed on the inner gap 16
  • b denotes a negative step formed on the inner gap 16.
  • a positive stage at the outer gap 17 is denoted by c
  • d denotes a negative step at the outer gap 17.
  • a positive stage a, c is present when the respective downstream wall 1 1, 12 protrudes radially into the respective gas path 9, 13 with respect to the respective upstream wall 7, 8.
  • there is a negative stage b, d when the respective upstream wall 7, 8 projecting radially projecting into the respective gas path 9, 13 with respect to the respective downstream wall 1 1, 12.
  • the positive stage a can be realized, for example, that the combustion chamber inner wall 7 in the adjacent to the gap 4 and the inner gap 16 end portion 21 in the positive stage a relative to the in the circumferential direction on both sides of the positive Stage a adjacent areas radially offset inwardly.
  • the inner positive stage a can thus be e.g. be realized only by contouring the chamber inner wall 7 in the end portion 21.
  • the inner positive stage a can be realized in that the turbine inner wall 1 1 extends in the associated end section 23 in the region of the positive step a radially outwardly offset relative to the areas adjacent to the positive step a in the circumferential direction. In this way, the positive stage a can be realized in principle only by a corresponding contouring of the turbine inner wall 11 in the end portion 23.
  • This negative step b can be realized in that the burner inner wall 7 in which the inner gap 16 adjacent end portion in the region of the negative stage b relative to the circumferentially on both sides of the negative step b adjacent areas radially offset outwardly extends.
  • the negative step b can also be realized in that the turbine inner wall 1 1 extends radially inwardly in the region adjacent to the inner gap 16 in the region of negative steps b relative to the regions adjacent to the negative step b in the circumferential direction.
  • the negative stage b can be realized by a combination of the above measures.
  • the positive stage c at the outer gap 17 can be realized in that the turbine outer wall 12 in the adjoining the outer gap 17 end portion 24 in the positive stage c relative to the circumferentially on both sides of the positive stage c adjacent areas radially inwardly offset runs.
  • the negative step d be realized, for example, that the combustion chamber outer wall 8 in the adjacent to the outer gap 17 end portion 22 in the negative stage d relative to the circumferentially on both sides of the negative step d adjacent areas radially inwardly staggered.
  • the negative step d can be realized at the outer gap 17 that the turbine outer wall 12 in the adjoining the outer gap 17 end portion 24 in the negative stage d relative to the circumferentially on both sides of the negative step d adjacent areas offset radially outwards. It is clear that a combination of the two above measures is also preferably realized in order to form the respective negative step d at the outer gap 17.
  • Gas paths 9, 13 alternate convex and concave areas, but gradually merge into each other.
  • the arrangement of positive and negative stages a, b, c, d which varies in the circumferential direction, may be formed as a function of a cooling requirement which occurs during operation of the gas turbine plant 1 at the respective gap 4 or at the inner gap 16 and / or at the outer gap 17 and especially in the circumferential direction can vary. It is clear that this cooling demand curve is present in the gap 16, 17 at a stationary operating state of the gas turbine plant 1 is substantially stationary. In order to increase the cooling gas flow in a peripheral segment that has an increased cooling requirement, a negative step b, d can be provided in this area.
  • a distribution in the circumferential direction of pressure sides and suction sides of the guide vanes 15 of the first row of guide vanes 14 can be taken into account.
  • These pressure sides and suction sides alternate in the circumferential direction and result from the profiling of the guide vanes 15.
  • alternating pressure sides and suction sides influence the pressure in the respective gas path 9, 13 also in the counterflow direction and at least up to the gap 4 a corresponding consideration of this distribution of the pressure sides and suction sides in the design of the steps at the gap 4, their influence can be reduced accordingly or used to set the desired cooling gas distribution.
  • flow conditions may arise in its gas path 9, which generate flow velocities or varying pressures varying at least in the outlet region 6 in the circumferential direction.
  • this pressure distribution varying in the circumferential direction can be stationary in a stationary operating state of the gas turbine plant 1. Accordingly, here too, the influence of a pressure distribution generated in the gap 4 by the operation of the combustion chamber 2 can be reduced or utilized for the desired cooling gas distribution by suitable design of stages a to d.
  • the measures proposed according to the invention are characterized by the fact that a significant influence of the cooling gas flow in the gap 4 can be realized without appreciably increasing the surface of the combustion chamber 2 or the turbine 3 exposed to the hot working gases becomes.
  • An enlarged surface as it is realized for example by projecting into the gas path Strömungsleitieri simultaneously increases the cooling demand for the flow guide and is disadvantageous in this respect.

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une installation de turbine à gaz (1), en particulier pour une centrale électrique, comprenant une chambre de combustion (2) qui limite radialement un chemin de gaz (9) de la chambre de combustion, au moins dans une zone de sortie annulaire (6) comprenant une paroi de chambre intérieure (7) et une paroi de chambre extérieure (8), une turbine (3) qui limite radialement un chemin de gaz (13) de turbine au moins dans une partie d'entrée (10) annulaire stationnaire comprenant une paroi de turbine intérieure (11) et une paroi de turbine extérieure (12), une fente (4; 16, 17) située radialement à l'intérieur et/ou radialement à l'extérieur et axialement entre la chambre de combustion (2) et la turbine (3), au niveau de laquelle se terminent la paroi de chambre intérieure et/ou extérieure (7, 8) et la paroi de turbine intérieure et/ou extérieure (11, 12), et un guidage de gaz de refroidissement (5) qui introduit par le biais de la fente (4; 16, 17) un gaz de refroidissement dans le chemin de gaz (13) de turbine et/ou dans le chemin de gaz (9) de la chambre de combustion. Afin d'augmenter le rendement de l'installation de turbine à gaz (1), on forme radialement à l'intérieur et/ou radialement à l'extérieur une portion d'extrémité (21, 22, 23, 24) adjacente à la fente (4; 16, 17) de la paroi de turbine respective (11, 12) et/ou de la paroi de chambre respective (7, 8) de telle sorte que des échelons positifs (a, c) dans le sens périphérique au niveau de la fente (4; 16, 17), alternent avec des échelons négatifs (b, d). Dans les échelons positifs, la paroi (11, 12) située en aval fait saillie radialement par rapport à la paroi respective située en amont (7, 8) en pénétrant dans le chemin de gaz respectif (9, 13). Dans les échelons négatifs, la paroi respective située en amont (7, 8) fait saillie radialement par rapport à la paroi respective située en aval (11, 12) en pénétrant dans le chemin de gaz respectif (9, 13).
EP08786927A 2007-08-06 2008-08-06 Refroidissement de fente entre une paroi de chambre de combustion et une paroi de turbine d'une installation de turbine à gaz Active EP2179143B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102007037070 2007-08-06
PCT/EP2008/060320 WO2009019282A2 (fr) 2007-08-06 2008-08-06 Installation de turbine à gaz

Publications (2)

Publication Number Publication Date
EP2179143A2 true EP2179143A2 (fr) 2010-04-28
EP2179143B1 EP2179143B1 (fr) 2011-01-26

Family

ID=40341811

Family Applications (1)

Application Number Title Priority Date Filing Date
EP08786927A Active EP2179143B1 (fr) 2007-08-06 2008-08-06 Refroidissement de fente entre une paroi de chambre de combustion et une paroi de turbine d'une installation de turbine à gaz

Country Status (5)

Country Link
US (1) US8132417B2 (fr)
EP (1) EP2179143B1 (fr)
AT (1) ATE497087T1 (fr)
DE (1) DE502008002497D1 (fr)
WO (1) WO2009019282A2 (fr)

Families Citing this family (13)

* 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
EP2248996B1 (fr) 2009-05-04 2014-01-01 Alstom Technology Ltd Turbine à gaz
CH703105A1 (de) 2010-05-05 2011-11-15 Alstom Technology Ltd Gasturbine mit einer sekundärbrennkammer.
EP2428647B1 (fr) 2010-09-08 2018-07-11 Ansaldo Energia IP UK Limited Zone de transition pour une chambre de combustion d'une turbine à gaz
DE102011008812A1 (de) 2011-01-19 2012-07-19 Mtu Aero Engines Gmbh Zwischengehäuse
US20160290645A1 (en) * 2013-11-21 2016-10-06 United Technologies Corporation Axisymmetric offset of three-dimensional contoured endwalls
US20170022839A1 (en) * 2013-12-09 2017-01-26 United Technologies Corporation Gas turbine engine component mateface surfaces
ES2632613T3 (es) 2014-08-29 2017-09-14 MTU Aero Engines AG Grupo constructivo de turbina de gas
DE102014221783A1 (de) * 2014-10-27 2016-04-28 Siemens Aktiengesellschaft Heißgaskanal
DE102014225689A1 (de) 2014-12-12 2016-07-14 MTU Aero Engines AG Strömungsmaschine mit Ringraumerweiterung und Schaufel
US10641174B2 (en) 2017-01-18 2020-05-05 General Electric Company Rotor shaft cooling
KR101958109B1 (ko) * 2017-09-15 2019-03-13 두산중공업 주식회사 가스 터빈
US10396795B1 (en) 2018-03-20 2019-08-27 Micron Technology, Inc. Boosted high-speed level shifter

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DE3023466C2 (de) * 1980-06-24 1982-11-25 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Einrichtung zur Verminderung von Sekundärströmungsverlusten in einem beschaufelten Strömungskanal
GB9304994D0 (en) * 1993-03-11 1993-04-28 Rolls Royce Plc Improvements in or relating to gas turbine engines
GB2281356B (en) * 1993-08-20 1997-01-29 Rolls Royce Plc Gas turbine engine turbine
EP0902164B1 (fr) * 1997-09-15 2003-04-02 ALSTOM (Switzerland) Ltd Refroidissement de la platte-forme dans les turbines à gas
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EP1515000B1 (fr) * 2003-09-09 2016-03-09 Alstom Technology Ltd Aubage d'une turbomachine avec un carenage contouré
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EP1731711A1 (fr) * 2005-06-10 2006-12-13 Siemens Aktiengesellschaft Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine
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Also Published As

Publication number Publication date
EP2179143B1 (fr) 2011-01-26
DE502008002497D1 (de) 2011-03-10
US20100146988A1 (en) 2010-06-17
US8132417B2 (en) 2012-03-13
WO2009019282A2 (fr) 2009-02-12
WO2009019282A3 (fr) 2009-05-07
ATE497087T1 (de) 2011-02-15

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