EP2342427B1 - Support d'aubes statorique axialement segmenté d'une turbine à gaz - Google Patents

Support d'aubes statorique axialement segmenté d'une turbine à gaz Download PDF

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Publication number
EP2342427B1
EP2342427B1 EP09824439.5A EP09824439A EP2342427B1 EP 2342427 B1 EP2342427 B1 EP 2342427B1 EP 09824439 A EP09824439 A EP 09824439A EP 2342427 B1 EP2342427 B1 EP 2342427B1
Authority
EP
European Patent Office
Prior art keywords
turbine
gas turbine
guide
blade carrier
axial
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.)
Not-in-force
Application number
EP09824439.5A
Other languages
German (de)
English (en)
Other versions
EP2342427A1 (fr
Inventor
Roderich Bryk
Sascha Dungs
Nicolas Savilius
Martin Hartmann
Uwe Kahlstorf
Karl Klein
Oliver Lüsebrink
Mirko Milazar
Oliver Schneider
Shilun Sheng
Vadim Shevchenko
Gerhard Simon
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 EP09824439.5A priority Critical patent/EP2342427B1/fr
Priority to PL09824439T priority patent/PL2342427T3/pl
Publication of EP2342427A1 publication Critical patent/EP2342427A1/fr
Application granted granted Critical
Publication of EP2342427B1 publication Critical patent/EP2342427B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F01D25/246—Fastening of diaphragms or stator-rings

Definitions

  • the invention relates to a guide vane carrier, in particular for a gas turbine, which consists of a number of axial segments.
  • 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.
  • For guiding the flow of the working medium in the turbine unit also commonly associated between adjacent blade rows with the turbine housing and combined into rows of guide vanes are arranged.
  • the combustion chamber of the gas turbine may be embodied as a so-called annular combustion chamber, in which a plurality of circumferentially arranged around the turbine shaft burners in a common, surrounded by a high temperature resistant surrounding wall combustion chamber space.
  • the combustion chamber is designed in its entirety as an annular structure.
  • a single combustion chamber can also be provided a plurality of combustion chambers.
  • first row of guide vanes of a turbine unit which, together with the blade row immediately downstream in the flow direction of the working medium, forms a first turbine stage of the turbine unit, which is usually followed by further turbine stages.
  • the vanes are fixed in each case via a blade root, also referred to as a platform, on a guide vane carrier of the turbine unit.
  • the guide blade carrier for securing the platforms of the guide vanes comprise an insulation segment.
  • a guide ring on the guide vane support of the turbine unit is arranged in each case.
  • Such a guide ring is spaced by a radial gap of the blade tips of the fixed at the same axial position on the turbine shaft blades of the associated blade row.
  • the guide vane carrier of the gas turbine usually made of cast steel. This is suitable to withstand the high temperatures within the gas turbine and it can thus ensure safe operation of the gas turbine.
  • the vanes of the gas turbine can be attached either to a common vane carrier or there are separate axial segments for each turbine stage, for example as in the GB 1 051 244 A , intended.
  • a common vane carrier or there are separate axial segments for each turbine stage, for example as in the GB 1 051 244 A , intended.
  • the turbine vane carrier is exposed to the extremely high temperatures that require high heat resistant cast steel, but there is a temperature profile that has relatively small high temperature regions and a larger, low temperature rear region.
  • the invention is therefore based on the object to provide a guide vane, which allows a technically simpler design and more flexible adaptation to the prevailing at the vane carrier temperature profile while maintaining operational safety.
  • This object is achieved according to the invention by designing at least one axial segment as a lattice tube structure.
  • the invention is based on the consideration that a more flexible adaptation to the temperature profile within the gas turbine could occur in the area of the guide blade carrier, in particular by different materials of the individual axial segments of the guide blade carrier.
  • high temperatures occur in particular in the region of the entanglement of the guide vanes and the ring segments, since these components ver a local heat input in the region of their attachment causes.
  • the foremost region of the guide blade carrier is exposed to a comparatively high compressor end temperature.
  • a relatively high quality material is necessary.
  • the temperature resistance of this material is not required. These areas could consist of cheaper and less expensive material.
  • the axial segments should continue to be solid in areas of low temperature. Therefore, these axial segments should be formed as a lattice structure with a plurality of tubes, rods, bars, beams, profiles or the like, ie as interconnected, arranged in the manner of a lattice structure struts.
  • the respective lattice structure is provided on its inner and / or outer side with a sheet metal lining.
  • a sheet metal lining for a special simple construction of the guide vane carrier is possible.
  • the embodiment with a sheet-metal-clad tubular construction can replace previously provided as castings sections of the vane support by a simpler structure, without jeopardizing the operational safety of the gas turbine. At the same time a smaller amount of material is needed.
  • the respective metal cladding on cooling air holes are also easier to manufacture than the cooling air holes required for castings, whereby a finer distribution to the subsequent ring segments can be provided by increasing the number of holes with the same cross-section or flow resistance.
  • the material of the respective axial segment and / or optionally of the respective Sheet metal cladding adapted to local thermal and mechanical loads during operation.
  • a number of axial segments are welded together.
  • the individual axial segments d. H. the individual lattice tube structures and the axial segments produced as castings a dimensionally stable and secure connection is ensured.
  • all axial segments are designed as a lattice tube structure.
  • the entire vane carrier can be designed as a lattice tube structure, where appropriate, different sheet metal linings are used on the inside, segment by segment. As a result, an even simpler construction of the guide vane carrier and thus the gas turbine is possible.
  • the advantages associated with the invention are, in particular, that a technically much simpler, easier and less expensive construction of a guide blade carrier and thus the entire gas turbine is possible by the design of an axial segment of a vane carrier as a lattice tube structure.
  • more favorable materials can be used in areas with lower temperature exposure and cost-intensive high-temperature materials remain restricted to the front, hot area of the gas turbine.
  • the remaining axial segments made of castings are comparatively smaller, allowing a simpler design of the vane carrier and the entire gas turbine.
  • the grid structure is less thermally conductive than a solid casting, also finds a lower heat conduction in the axial direction, in particular from the hot areas at the compressor exit in the rear cooler areas, resulting in improved cooling of the vane support and thus a lower axial and possibly also radial thermal Expansion is achieved.
  • this design shows great potential for further development of guide vane carrier, as more flexible can be addressed to thermal and mechanical requirements.
  • At the front of the turbine vane carrier there are extremely stringent requirements for maintaining the gap to the vanes and blades to ensure turbine efficiency. With the segmentation by the tubular structure, the thermal expansion behavior can be set to a much better extent than before, and thus the necessary minimum gap can be reduced.
  • FIG. 1 shows in detail a half section through a guide vane carrier 1.
  • Leit Blade carrier 1 In stationary gas turbines is the Leit Blade carrier 1 usually conical or cylindrical shaped and consists of two segments, an upper and a lower segment, the z. B. are interconnected via flanges. Only the section through the upper segment is shown.
  • the illustrated vane carrier 1 comprises a number of axial segments 24 which are welded together to form a solid structure.
  • the grid structures 26 are each provided on their inner side with a sheet metal lining 28.
  • the struts of the grid construction can be designed with a variety of profiles such as round, square or otherwise as a hollow body or in solid construction.
  • the remaining axial segments 24 are formed as castings 30.
  • the material of the cast parts 30 and the sheet metal linings 28 is in each case adapted to the thermal conditions in their respective region in the interior of the gas turbine.
  • a complete construction of the vane support 1 made of grid segments would also be possible.
  • the gas turbine 101 includes 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 a common turbine shaft 108, also referred to as a turbine runner, to which the generator or the working machine is also connected and which is rotatably mounted about its central axis 109.
  • the running in the manner of an annular combustion chamber combustion chamber 104 is provided with a number of Burners 110 equipped for the combustion of a liquid or gaseous fuel.
  • the turbine unit 106 has a number of rotatable blades 112 connected to the turbine shaft 108.
  • the blades 112 are annularly disposed on the turbine shaft 108 and thus form a number of blade rows.
  • the turbine unit 106 includes a number of stationary vanes 114, which are also attached in a donut-like manner to a vane support 1 of the turbine unit 106 to form rows of vanes.
  • the blades 112 serve to drive the turbine shaft 108 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 112 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 guide ring 121 is arranged on the guide blade carrier 1 of the turbine unit 106.
  • the outer surface of each guide ring 121 is also the hot, the turbine unit 106 flowing through the working medium M and radially spaced from the outer end of the opposed blades 112 by a gap.
  • the guide rings 121 arranged between adjacent rows of guide blades serve in particular as cover elements which protect the inner housing in the guide blade carrier 1 or other housing built-in components against thermal overstress by the hot working medium M flowing through the turbine 106.
  • 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 108 around.
  • the leftmost axial segments 24 are accordingly made of a high temperature resistant material than in the gas channel downstream areas.
  • the lattice structure furthermore ensures good thermal insulation of the individual cast parts 30 from one another, as a result of which thermal deformations can be minimized.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (8)

  1. Support ( 1 ) d'aube directrice pour une turbine ( 101 ) à gaz, qui est constituée d'un certain nombres de segments ( 24 ) axiaux, caractérisé en ce qu'au moins un segment ( 24 ) axial est conçu sous la forme d'une structure ( 26 ) tubulaire grillagée.
  2. Support ( 1 ) d'aube directrice suivant la revendication 1, dans lequel la structure ( 26 ) tubulaire grillagée est pourvue d'un habillage ( 28 ) de tôle sur son côté intérieur et/ou sur son côté extérieur.
  3. Support ( 1 ) d'aube directrice suivant la revendication 2, dans lequel l'habillage ( 28 ) de tôle a des trous d'air de refroidissement.
  4. Support ( 1 ) d'aube directrice suivant l'une des revendications 1 à 3, dans lequel le matériau du segment ( 24 ) axial et/ou le cas échéant de l'habillage ( 28 ) de tôle est adapté aux charges locales thermiques et mécaniques prévues en fonctionnement.
  5. Support ( 1 ) d'aube directrice suivant l'une des revendications 1 à 4, dans lequel un certain nombres de segments ( 24 ) axiaux sont soudés les uns aux autres.
  6. Support ( 1 ) d'aube directrice suivant l'une des revendications 1 à 5, dans lequel tous les segments ( 24 ) axiaux sont conçus sous la forme d'une structure ( 28 ) tubulaire grillagée.
  7. Turbine ( 101 ) à gaz ayant un support ( 1 ) d'aube directrice suivant l'une des revendications 1 à 6.
  8. Installation de turbine à gaz et de turbine à vapeur ayant une turbine ( 101 ) à gaz suivant la revendication 7.
EP09824439.5A 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz Not-in-force EP2342427B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP09824439.5A EP2342427B1 (fr) 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz
PL09824439T PL2342427T3 (pl) 2008-11-05 2009-09-10 Dźwigar łopatek kierujących złożony z segmentów osiowych dla turbiny gazowej

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP08019365A EP2184445A1 (fr) 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz
EP09824439.5A EP2342427B1 (fr) 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz
PCT/EP2009/061744 WO2010052050A1 (fr) 2008-11-05 2009-09-10 Support d'aubes fixes, segmenté axialement, pour une turbine à gaz

Publications (2)

Publication Number Publication Date
EP2342427A1 EP2342427A1 (fr) 2011-07-13
EP2342427B1 true EP2342427B1 (fr) 2013-06-19

Family

ID=40497476

Family Applications (2)

Application Number Title Priority Date Filing Date
EP08019365A Withdrawn EP2184445A1 (fr) 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz
EP09824439.5A Not-in-force EP2342427B1 (fr) 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP08019365A Withdrawn EP2184445A1 (fr) 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz

Country Status (7)

Country Link
US (1) US8870526B2 (fr)
EP (2) EP2184445A1 (fr)
JP (1) JP5596042B2 (fr)
CN (1) CN102216568B (fr)
PL (1) PL2342427T3 (fr)
RU (1) RU2508450C2 (fr)
WO (1) WO2010052050A1 (fr)

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US10094287B2 (en) 2015-02-10 2018-10-09 United Technologies Corporation Gas turbine engine component with vascular cooling scheme
US10077664B2 (en) 2015-12-07 2018-09-18 United Technologies Corporation Gas turbine engine component having engineered vascular structure
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Also Published As

Publication number Publication date
WO2010052050A1 (fr) 2010-05-14
RU2011122612A (ru) 2012-12-20
PL2342427T3 (pl) 2013-11-29
EP2184445A1 (fr) 2010-05-12
JP2012507652A (ja) 2012-03-29
US8870526B2 (en) 2014-10-28
CN102216568A (zh) 2011-10-12
RU2508450C2 (ru) 2014-02-27
EP2342427A1 (fr) 2011-07-13
JP5596042B2 (ja) 2014-09-24
US20110268580A1 (en) 2011-11-03
CN102216568B (zh) 2015-11-25

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