EP2218880A1 - Système de contrôle actif de jeu pour turbine à gaz - Google Patents

Système de contrôle actif de jeu pour turbine à gaz Download PDF

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Publication number
EP2218880A1
EP2218880A1 EP09002137A EP09002137A EP2218880A1 EP 2218880 A1 EP2218880 A1 EP 2218880A1 EP 09002137 A EP09002137 A EP 09002137A EP 09002137 A EP09002137 A EP 09002137A EP 2218880 A1 EP2218880 A1 EP 2218880A1
Authority
EP
European Patent Office
Prior art keywords
membrane
guide
turbine
guide vane
gas turbine
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.)
Withdrawn
Application number
EP09002137A
Other languages
German (de)
English (en)
Inventor
François Dr. Benkler
Tobias Dr. Buchal
Andreas Dr. Böttcher
Martin Hartmann
Patricia Dr. Hülsmeier
Dieter Minninger
Oliver Dr. Schneider
Norbert Thamm
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.)
Siemens AG
Original Assignee
Siemens 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 Siemens AG filed Critical Siemens AG
Priority to EP09002137A priority Critical patent/EP2218880A1/fr
Publication of EP2218880A1 publication Critical patent/EP2218880A1/fr
Withdrawn legal-status Critical Current

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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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/22—Actively adjusting tip-clearance by mechanically actuating the stator or rotor components, e.g. moving shroud sections relative to the rotor
    • 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/207—Heat transfer, e.g. cooling using a phase changing mass, e.g. heat absorbing by melting or boiling
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/232—Heat transfer, e.g. cooling characterized by the cooling medium

Definitions

  • the invention relates to a vane carrier system, in particular for a gas turbine, with a number of vanes combined into a vane carrier and attached to a vane carrier. It further relates to a gas turbine with such a Leitschaufelitatisystem.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • a number of rotor blades which are usually combined into blade groups or rows of blades, are arranged thereon and drive the turbine shaft via a momentum transfer from the working medium.
  • guide vanes are also usually arranged between adjacent rotor blade rows and connected to the turbine housing, which are combined into rows of guide blades. These are attached to a usually hollow cylindrical or hollow cone-shaped vane carrier.
  • the hot gas channel is usually lined by so-called ring segments, which form the inner wall of the hot gas channel. These are usually fastened via hooking elements on the guide blade carrier, so that the entirety of the ring segments in the circumferential direction as well as the guide blade carrier form a hollow conical or hollow cylindrical structure.
  • the components of the gas turbine can be deformed by different thermal expansion in different operating conditions, which has a direct influence on the size of the radial gaps between the blades and the inner wall of the hot gas duct, d. H. has the ring segments.
  • These radial gaps are differently dimensioned when starting and stopping the turbine than in regular operation.
  • inner wall or blades are always to be dimensioned so that the radial gaps are kept sufficiently large to cause damage to the gas turbine in any operating condition.
  • a correspondingly comparatively generous design of the radial gaps leads to considerable losses in the efficiency.
  • the invention is therefore based on the object to provide a guide vane system, which allows a particularly high efficiency while maintaining the greatest possible operational safety and durability.
  • This object is achieved according to the invention by arranging a membrane which can be acted upon by a pressure medium in one axially adjacent region of the respective row of guide blades.
  • the invention is based on the consideration that a particularly high efficiency by reducing the radial gap in normal operation, d. H.
  • a particularly high efficiency by reducing the radial gap in normal operation d. H.
  • a comparatively large dimension of the radial gaps is required in particular because the components of the gas turbine deform differently in different operating states.
  • other radial gaps are present than in regular operation.
  • set in partial load operation also different column than in full load operation.
  • responsible for this temporal change of the radial gap are the different thermal inertia behavior of the individual components, the centrifugal force expansion and transverse contraction of the rotor, the play in the thrust bearing and the ovalization of the housing as a result of montage employmenter bias and uneven heating.
  • an adaptation of the radial gaps should not be done by a corresponding design in the construction of the gas turbine or the Leitschaufelanisystems, but it should be adaptive adjustment of the radial gaps during operation of the gas turbine .
  • the radial gaps are determined by the distance between the blade tips to the respective opposite, the guide vanes adjacent areas of the guide blade carrier. Therefore, an adaptive adaptation of the radial gaps could be achieved by a possibility for radial movement of the inner wall components of the guide vane carrier.
  • a membrane which can be acted upon by a pressure medium, i. H. be arranged an elastic, stretchable surface.
  • the membrane is arranged in a region of the guide blade carrier which is intended for enclosing a blade row of the gas turbine.
  • a particularly simple construction of such a membrane is possible if the membrane forms a peripheral wall of a chamber which can be acted upon by the pressure medium.
  • the chamber then forms a pressure chamber, which can be acted upon by the pressure medium in each required amount.
  • the membrane should form each of the turbine axis facing the surrounding wall of the chamber.
  • the elastic membrane then expands more or less as the surrounding wall of the chamber and thus influences the extent of the radial gaps.
  • ring segments are arranged in the regions adjacent to the guide vanes, which segments form the inner peripheral wall of the hot gas duct of the gas turbine and protect the guide vanes from damage due to the penetration of hot working medium. Therefore, in an advantageous embodiment, the pressurizable with the pressure medium chamber should be arranged in such, attached to the guide vane ring segment. This allows an adaptive reduction of the radial gaps, while the protection of the guide vane carrier against ingress of hot gas remains guaranteed.
  • channels are introduced into the guide vane carrier system for feeding the pressure medium to the membrane.
  • These may be incorporated, for example, in the vane carrier and the ring segments, which allows easy delivery and control of the pressure on the membrane from the outside during operation of the gas turbine.
  • the radial extent of the membrane should be able to be influenced separately depending on the circumferential position.
  • a plurality of membranes and possibly chambers is advantageously arranged in the circumferential direction, so that a separate admission of the membranes or chambers with different pressures is possible.
  • a greater pressure can be deliberately applied to the membrane in peripheral areas with larger radial gaps that are set up so that a uniform, comparatively small radial gap is achieved over the entire circumference.
  • the pressure medium is sodium.
  • Sodium has particularly good heat transfer properties, a low melting point and at the same time a large liquid temperature range and is thus particularly suitable as a pressure medium.
  • such a vane carrier is part of a gas turbine and such a gas turbine part of a combined cycle power plant.
  • the advantages achieved by the invention are, in particular, that an adaptive hydraulic adjustment of the radial gaps during operation is made possible by the arrangement of an elastic membrane which can be acted upon by a pressure medium in an area of a guide blade carrier adjacent to the respective row of guide blades.
  • the membrane When the membrane is pressurized with the pressure medium, the membrane is expanded towards the blade and thus the ring diameter is reduced. This also reduces the radial gap between the blade and the housing. This makes it possible to achieve an improvement in the efficiency of the gas turbine, with the adaptation of the radial gaps only taking place during operation and not necessarily influencing the design in the construction of the gas turbine.
  • Such a hydraulic radial gap closure also eliminates the disadvantages of previous solutions such as an axial displacement of the entire turbine shaft to the compressor inlet, which causes a reduction of the radial column in the turbine by the conical shape of the vane support, but an increase in the gap and thus a reduction in efficiency in the compressor result.
  • the reduction of the radial gaps by a membrane which can be acted upon by a pressure medium thus enables a targeted individual radial gap adjustment for compressor and turbine and thus enables a particularly high efficiency of the entire gas turbine.
  • FIG. 1 shows a part of a Leitschaufelitationsystems 1, here in particular one on a guide vane not shown here fixed ring segment 2, which forms the inner wall 4 of the hot gas channel of a gas turbine and adjacent to a number of blades not shown in detail surrounds a number of blades 6.
  • the rotor blades 6 are arranged on a turbine shaft 8 and rotate during operation of the gas turbine.
  • the radial gap 10 between the tips 12 of the guide vanes 8 and the inner wall 4 should be kept as low as possible.
  • a number of chambers 14 is introduced into the ring segment 2, the turbine axis facing the surrounding wall is formed from a membrane 16.
  • the membrane 16 is in the FIG. 1 shown in the relaxed state and can expand when exposed to a corresponding pressure medium.
  • 2 channels 18 are introduced into the ring segment, through which the chamber 14 with a pressure medium, such as sodium, can be filled. This leads to a corresponding admission of the membrane 16.
  • a plurality of separate chambers 14 and diaphragms 16 is provided, which makes it possible, depending on the circumferential position, to apply different pressures to the chambers 14 in order to control the radial extent of the respective membrane 16 separately. This makes it possible in particular to compensate for ovalization of the ring segments 2 caused by thermal deformation of the guide blade carrier.
  • the guide vane support system 1 shown with the ring segment 2 can be used both within the turbine of a gas turbine and in the compressor.
  • the 3 and 4 show once again the representations of FIG. 1 respectively.
  • FIG. 2 in a different perspective.
  • a gas turbine 101 as in FIG. 5 has a compressor 102 for combustion air, a combustion chamber 104 and a turbine unit 106 for driving the compressor 102 and a generator, not shown, or a working machine.
  • the turbine unit 106 and the compressor 102 are arranged on the common turbine shaft 8, which is also referred to as a turbine runner and to which the generator or the work machine is also connected, and which is mounted rotatably about its turbine axis 109.
  • the combustor 104 which is in the form of an annular combustor, is equipped with a number of burners 110 for combustion of a liquid or gaseous fuel.
  • the turbine unit 106 has a number of rotatable blades 6 connected to the turbine shaft 108.
  • the blades 6 are arranged in a ring shape on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine unit 106 includes a number of stationary vanes 114 which are also annularly attached to a vane support 116 of the turbine unit 106 to form rows of vanes.
  • the blades 6 serve to drive the turbine shaft 8 by momentum transfer from the turbine unit 106 flowing through the working medium M.
  • the vanes 114 serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 114 or a row of vanes and a ring of blades 6 or a blade row is also referred to as a turbine stage.
  • Each vane 114 has a platform 118 which is arranged to fix the respective vane 114 to a vane support 1 of the turbine unit 106 as a wall element.
  • the platform 118 is a thermally comparatively heavily loaded component, which forms the outer boundary of a hot gas channel for the turbine unit 106 flowing through the working medium M.
  • Each blade 112 is fastened to the turbine shaft 108 in an analogous manner via a platform 119, also referred to as a blade root.
  • a ring segment 2 is arranged on a guide blade carrier 116 of the turbine unit 106.
  • the outer surface of each ring segment 2 is also exposed to the hot, the turbine unit 106 flowing through the working fluid M and spaced in the radial direction from the outer end of the opposite blades 6 through the radial gap 10.
  • the arranged between adjacent rows of stator ring segments 2 are used in particular as cover that protect the inner housing in the guide blade carrier 1 or other housing-mounting components from thermal overload by the turbine 106 flowing through the hot working medium M.
  • the combustion chamber 104 is configured in the exemplary embodiment as a so-called annular combustion chamber, in which a plurality of burners 110 arranged around the turbine shaft 108 in the circumferential direction open into a common combustion chamber space.
  • the combustion chamber 104 is configured in its entirety as an annular structure, which is positioned around the turbine shaft 8 around.
  • a particularly high efficiency can be achieved with simultaneously high operational safety and service life. Due to the hydraulic radial gap optimization, the radial gaps 10 can be separately optimized in each operating state both in the compressor 102 and in the turbine unit 106 of the gas turbine 101.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP09002137A 2009-02-16 2009-02-16 Système de contrôle actif de jeu pour turbine à gaz Withdrawn EP2218880A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP09002137A EP2218880A1 (fr) 2009-02-16 2009-02-16 Système de contrôle actif de jeu pour turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09002137A EP2218880A1 (fr) 2009-02-16 2009-02-16 Système de contrôle actif de jeu pour turbine à gaz

Publications (1)

Publication Number Publication Date
EP2218880A1 true EP2218880A1 (fr) 2010-08-18

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EP09002137A Withdrawn EP2218880A1 (fr) 2009-02-16 2009-02-16 Système de contrôle actif de jeu pour turbine à gaz

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9039346B2 (en) 2011-10-17 2015-05-26 General Electric Company Rotor support thermal control system

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2458676A1 (fr) * 1979-06-06 1981-01-02 Mtu Muenchen Gmbh Dispositif d'etancheite pour l'intervalle peripherique d'une turbomachine a flux axial
GB2050527A (en) * 1979-05-29 1981-01-07 Gen Motors Corp Turbine blade tip seal assembly
US4419044A (en) * 1980-12-18 1983-12-06 Rolls-Royce Limited Gas turbine engine
GB2195715A (en) * 1986-10-08 1988-04-13 Rolls Royce Plc Rotor blade tip-shroud
US5022817A (en) * 1989-09-12 1991-06-11 Allied-Signal Inc. Thermostatic control of turbine cooling air
US5344284A (en) * 1993-03-29 1994-09-06 The United States Of America As Represented By The Secretary Of The Air Force Adjustable clearance control for rotor blade tips in a gas turbine engine
DE102006052786A1 (de) * 2006-11-09 2008-05-15 Mtu Aero Engines Gmbh Turbomaschine

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2050527A (en) * 1979-05-29 1981-01-07 Gen Motors Corp Turbine blade tip seal assembly
FR2458676A1 (fr) * 1979-06-06 1981-01-02 Mtu Muenchen Gmbh Dispositif d'etancheite pour l'intervalle peripherique d'une turbomachine a flux axial
US4419044A (en) * 1980-12-18 1983-12-06 Rolls-Royce Limited Gas turbine engine
GB2195715A (en) * 1986-10-08 1988-04-13 Rolls Royce Plc Rotor blade tip-shroud
US5022817A (en) * 1989-09-12 1991-06-11 Allied-Signal Inc. Thermostatic control of turbine cooling air
US5344284A (en) * 1993-03-29 1994-09-06 The United States Of America As Represented By The Secretary Of The Air Force Adjustable clearance control for rotor blade tips in a gas turbine engine
DE102006052786A1 (de) * 2006-11-09 2008-05-15 Mtu Aero Engines Gmbh Turbomaschine

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9039346B2 (en) 2011-10-17 2015-05-26 General Electric Company Rotor support thermal control system

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