WO2010049195A1 - Turbine à gaz avec insert de refroidissement - Google Patents

Turbine à gaz avec insert de refroidissement Download PDF

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Publication number
WO2010049195A1
WO2010049195A1 PCT/EP2009/061461 EP2009061461W WO2010049195A1 WO 2010049195 A1 WO2010049195 A1 WO 2010049195A1 EP 2009061461 W EP2009061461 W EP 2009061461W WO 2010049195 A1 WO2010049195 A1 WO 2010049195A1
Authority
WO
WIPO (PCT)
Prior art keywords
cooling
cooling air
gas turbine
turbine
guide
Prior art date
Application number
PCT/EP2009/061461
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Christian Lerner
Original Assignee
Siemens Aktiengesellschaft
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 Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Priority to CN200980142781.5A priority Critical patent/CN102197195B/zh
Priority to EP09823098A priority patent/EP2347100B1/fr
Priority to JP2011532553A priority patent/JP5281166B2/ja
Priority to US13/125,978 priority patent/US20110255956A1/en
Publication of WO2010049195A1 publication Critical patent/WO2010049195A1/fr

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
    • 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
    • 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
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • the invention relates to a gas turbine having a number of rotor blades arranged in each case, arranged on a turbine shaft and with a number of guide vanes combined by vanes on a turbine housing fixed guide vanes, the guide vanes having a number of cooling air holes.
  • 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 which is 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 usually arranged between adjacent rotor blade rows and connected to the turbine housing and combined into rows of guide blades.
  • the combustion chamber of the gas turbine can be embodied as a so-called annular combustion chamber, in which a multiplicity of burners arranged around the turbine shaft in the circumferential direction can be divided into a common, high-temperature-resistant burner. gen surrounding wall combustion chamber space opens.
  • the combustion chamber is designed in its entirety as an annular structure.
  • a plurality of combustion chambers In addition to 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 guide vanes are each fixed to a vane support of the turbine unit via a blade root, also referred to as a platform.
  • 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 above-mentioned guide rings can, as known, for example, from US Pat. No. 3,864,056, be cooled. According to US 3,864,056, the guide ring segments are hooked to the vane carrier. In the wall is a supply of cooling air to the
  • Guide rings provided in the form of a passage.
  • a preloaded sleeve is screwed, which presses the guide ring segment against the hook, wherein the in Inside the sleeve flowing cooling air can pass through openings in the cold gas side rear space of the guide ring segment and is further used there for cooling the guide ring segment.
  • An alternative mounting and cooling of guide ring segments is shown in GB 1 524 956.
  • bores may be provided in the guide blade carrier, are guided by the measuring lances, with which the radial gap between the guide ring segment and blade tip is detected.
  • a cooled measuring lance is known from US 2006/0140754 A1.
  • the guide vane carrier of the gas turbine is usually made of cast steel, since this is suitable to withstand the high temperatures within the gas turbine.
  • cooling air bores are usually provided in the guide blade carrier, through which cooling air flows from the outer regions of the gas turbine into the interior, thereby cooling the guide blade carrier.
  • a plurality of cooling air reservoirs with different temperatures and pressures between the turbine housing and the guide blade carrier are provided.
  • Adequate cooling of the vane support is necessary, inter alia, because too high temperatures and thus too high temperature differences in different operating states cause the thermal deformations of the guide vane wearer, which must be considered in the construction of the gas turbine.
  • the gap dimensions in particular of the radial gaps between the blade ends and the inner wall, must be selected to be correspondingly large in order to compensate for variances produced by the deformation of the guide blade carrier and thus prevent damage to the gas turbine.
  • Increasing the gap results in a reduction in the efficiency of the gas turbine. Accordingly, sufficient cooling should always be provided to reduce deformation of the vane support.
  • a strong cooling of the vane carrier also means a high consumption of cooling air, which then flows into the interior of the gas turbine. This lowers the temperature in the
  • the invention is therefore based on the object to provide a gas turbine, which has a particularly high efficiency while maintaining the greatest possible operational safety.
  • a cooling insert is introduced into at least one cooling air hole to the wall cooling.
  • the invention is based on the consideration that a particularly high efficiency can be achieved by increasing the temperature inside the gas turbine. This can be done by reducing the cooling air consumption, ie a reduction in the amount of cooling air introduced into the interior of the gas turbine.
  • a reduction in the amount of cooling air results in an increase in the temperature of the vane carrier, since less air then flows through its cooling air bores and accordingly less heat is removed from the vane carrier.
  • this can result in deformation of the vane carrier, the in the construction of the gas turbine then would have to be considered. Therefore, the existing cooling air should be used very effectively for cooling, ie, it should be removed with the lowest possible amount of cooling air, the largest possible amount of heat.
  • the cooling insert is tubular and provided with arranged in its tube wall, window-shaped wall openings. This makes it possible that the cooling air flowing through the cooling insert can continue to come into contact with the wall of the cooling air holes of the guide blade carrier in order to remove the heat energy.
  • the wall openings are a large area and separated by webs, whereby the cooling air over a large area can come into contact with the wall of the cooling air hole.
  • the respective cooling insert comprises at least one turbulator.
  • Turbulators are small surveys, ie generally applied surface disturbances, which convert a laminar flow into a turbulent one. These can be formed, for example, by the webs or in the form of raised wires, sheet corners or the like. Even if the flow in the cooling air bore is already turbulent, these turbulators ensure even better heat transfer and thus overall better cooling of the guide blade carrier with reduced cooling air consumption.
  • the cooling insert can advantageously also be designed as an impingement cooling insert, for example when the wall openings are designed as impingement cooling openings which are distributed like a grid. The cooling air flowing through the cooling insert can radiate out through the impingement cooling apertures while impinging transversely on the cooling air bore walls of the guide vane support. As a result, a particularly efficient cooling of the guide blade carrier is achieved.
  • the respective cooling insert comprises thread-like structures.
  • a thread structure of the flow inside the cooling air hole By a thread structure of the flow inside the cooling air hole, a twist can be imposed, which on the one hand ensures a turbulence of the flow, on the other hand has a longer whereabouts of the cooling air in the cooling air bore. As a result, a better heat transfer from the material of the vane support is ensured on the flowing cooling air also.
  • the respective cooling insert is made of the same material as the guide blade carrier. This may possibly complications due to different choice of material of the cooling insert and the Leitschaufelträ- gers, such as a different thermal
  • cooling air supply line to the cooling air bores should advantageously be adapted to the cooling properties of the respective cooling insert.
  • the temperature and pressure of the introduced cooling air are optimized to the new, changed properties with regard to the cooling by the cooling inserts.
  • a gas turbine is used in a gas and steam turbine plant.
  • the advantages associated with the invention are, in particular, that an overall better efficiency of the gas turbine is achieved by the introduction of cooling inserts in the cooling air holes of the guide blade carrier by the improved cooling while reducing the amount of cooling air. Furthermore, such inserts can be particularly easy to introduce and can therefore also be relatively easily used in the manner of retrofitting in older gas turbines.
  • the cooling inserts can also be flexibly adapted to the respective requirements with regard to cooling and cooling air consumption.
  • FIG 3 shows a plan view of a cooling insert.
  • the gas turbine 1 according to FIG. 1 has a compressor 2 for
  • Combustion air a combustion chamber 4 and a turbine unit 6 for driving the compressor 2 and a generator, not shown, or a working machine on.
  • the turbine unit 6 and the compressor 2 are arranged on a common turbine shaft 8, also referred to as a turbine rotor, to which the generator or the working machine is also connected, and which is rotatably mounted about its central axis 9.
  • the type of an annular combustion chamber te combustion chamber 4 is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel.
  • the turbine unit 6 has a number of rotatable blades 12 connected to the turbine shaft 8.
  • the blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine unit 6 comprises a number of fixed guide vanes 14, which are also fixed in a ring shape with the formation of rows of vanes on a guide vane support 16 of the turbine unit 6.
  • the blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine unit 6 flowing through the working medium M.
  • a successive pair of a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.
  • Each vane 14 has a platform 18 which is arranged to fix the respective vane 14 to a vane support 16 of the turbine unit 6 as a wall element.
  • the platform 18 is a thermally comparatively heavily loaded component which forms the outer boundary of a hot gas channel for the working medium M flowing through the turbine unit 6.
  • Each rotor blade 12 is fastened to the turbine shaft 8 in an analogous manner via a platform 19, also referred to as a blade root.
  • each guide ring 21 is arranged on a guide blade carrier 16 of the turbine unit 6.
  • the outer surface of each guide ring 21 is also the hot, exposed to the turbine unit 6 flowing through the working medium M and spaced in the radial direction from the outer end of the opposed blades 12 by a gap.
  • the guide rings 21 arranged between adjacent guide blade rows serve in particular as cover elements which protect the inner housing 16 in the guide blade carrier or other housing built-in components against thermal overstress by the hot working medium M flowing through the turbine 6.
  • the combustion chamber 4 is designed in the embodiment as a so-called annular combustion chamber, in which a plurality of circumferentially around the turbine shaft 8 arranged around burners 10 open into a common combustion chamber space.
  • the combustion chamber 4 in its entirety as an annular
  • Structured structure which is positioned around the turbine shaft 8 around.
  • cooling air bores are introduced into the guide blade carrier 16 by the cooling air of different temperature and pressure from different chambers outside the region of the guide blade carrier 16 through the guide vane carrier 16 into the interior of the gas turbine 1 is performed. This cooling air ensures cooling of the guide blade carrier 16 so that thermal deformations of the guide blade carrier 16 are reduced.
  • cooling inserts 22 are inserted into the cooling air bores. If the cooling insert 22 is designed as an impact-cooling insert, its outer diameter is slightly smaller than the diameter of the cooling-air bore. A cross section through one half of such a cooling insert 22 is shown in FIG.
  • the cooling insert 22 has a substantially cylindrical shape in order to be used in the existing cooling air holes can. To this
  • cooling insert 22 is tubular, so it can flow through along its axial extent.
  • flange 23 On one side of the cooling insert 22 includes a flange 23 for fixing.
  • the cooling insert 22 has at its circular cross-section pipe wall a plurality of window-shaped wall openings 25, which can be distributed both along its axial extent and on the circumference.
  • the wall openings are comparatively large area and are separated by webs 26 from each other.
  • Such a cooling insert 22 then has, in contrast to the impingement cooling insert, an outer diameter which corresponds to the diameter of the cooling air bore.
  • the webs 26 extending in the circumferential direction of the cooling insert 22 are designed as turbulators 24, at which the air flow breaks and the laminar flow is transformed into a turbulent flow.
  • turbulators Other forms and arrangements of turbulators are also possible.
  • Turbulators 24 may also be arranged in the manner of a thread, so that the cooling air is given an additional twist, so that the dwell time and the turbulence in the cooling air hole is larger.
  • FIG. 3 shows the cooling air insert 22 once again in the plan view. Again, the flanges 23 for fixing in the cooling air holes of the vane support 16 can be seen. There is improved by the cooling insert 22, the heat transfer from the material of the vane support 16 to the cooling air in the cooling air holes, further, the cooling air supply to the vane support 16 should still be adapted to the new cooling air properties. As a result, a comparatively better and more effective cooling of the vane support 16 is ensured while at the same time reducing the consumption of cooling air. As a result, the efficiency of the gas turbine 1 can be increased overall.

Landscapes

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

Abstract

L'invention concerne une turbine à gaz (1) avec un certain nombre d'aubes mobiles (12) montées sur un arbre de turbine (8) et réunies en rangées respectives d'aubes mobiles et avec un certain nombre d'aubes fixes (14) fixées par un support (16) d'aubes fixes sur un carter de la turbine et réunies en rangées respectives d'aubes fixes. Le support (16) d'aubes fixes présente un certain nombre d'alésages d'air de refroidissement. L'objectif est d'atteindre un rendement particulièrement élevé en conservant la sécurité de fonctionnement la plus élevée possible. À cet effet, un insert de refroidissement (22) est introduit dans un alésage d'air de refroidissement.
PCT/EP2009/061461 2008-10-27 2009-09-04 Turbine à gaz avec insert de refroidissement WO2010049195A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
CN200980142781.5A CN102197195B (zh) 2008-10-27 2009-09-04 具有冷却嵌件的燃气涡轮机
EP09823098A EP2347100B1 (fr) 2008-10-27 2009-09-04 Turbine à gaz avec insert de refroidissement
JP2011532553A JP5281166B2 (ja) 2008-10-27 2009-09-04 冷却用インサートを有したガスタービン
US13/125,978 US20110255956A1 (en) 2008-10-27 2009-09-04 Gas turbine having cooling insert

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP08018754A EP2180148A1 (fr) 2008-10-27 2008-10-27 Turbine à gaz avec noyau de refroidissement
EP08018754.5 2008-10-27

Publications (1)

Publication Number Publication Date
WO2010049195A1 true WO2010049195A1 (fr) 2010-05-06

Family

ID=40474905

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2009/061461 WO2010049195A1 (fr) 2008-10-27 2009-09-04 Turbine à gaz avec insert de refroidissement

Country Status (5)

Country Link
US (1) US20110255956A1 (fr)
EP (2) EP2180148A1 (fr)
JP (1) JP5281166B2 (fr)
CN (1) CN102197195B (fr)
WO (1) WO2010049195A1 (fr)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103075202A (zh) * 2013-01-15 2013-05-01 上海交通大学 涡轮叶片内部带有栅格扰流的冲击冷却结构
EP3023600B1 (fr) * 2014-11-24 2018-01-03 Ansaldo Energia IP UK Limited Élément de carter de moteur
RU2740069C1 (ru) * 2017-12-01 2020-12-31 Сименс Энерджи, Инк. Впаянный теплопередающий элемент для охлаждаемых компонентов турбины
CN112177689A (zh) * 2020-09-29 2021-01-05 中国航发湖南动力机械研究所 发动机的涡轮导向器定位结构和发动机

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3864056A (en) * 1973-07-27 1975-02-04 Westinghouse Electric Corp Cooled turbine blade ring assembly
GB1524956A (en) * 1975-10-30 1978-09-13 Rolls Royce Gas tubine engine
US5253976A (en) * 1991-11-19 1993-10-19 General Electric Company Integrated steam and air cooling for combined cycle gas turbines
EP1028230A1 (fr) * 1999-02-09 2000-08-16 ABB Alstom Power (Schweiz) AG Pièce refroidie de turbine à gaz avec refroidissement ajustable
US20040219009A1 (en) * 2003-03-06 2004-11-04 Snecma Moteurs Turbomachine with cooled ring segments
EP1526251A1 (fr) * 2003-10-22 2005-04-27 General Electric Company Configuration de refroidissement pour une aube de turbine
US20060140754A1 (en) * 2004-12-27 2006-06-29 Mitsubishi Heavy Industries, Ltd. Gas turbine
EP1923538A2 (fr) * 2006-11-15 2008-05-21 General Electric Company Turbine avec réglage du jeu des extrémités des aubes par transpiration

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP4274666B2 (ja) * 2000-03-07 2009-06-10 三菱重工業株式会社 ガスタービン
US7056084B2 (en) * 2003-05-20 2006-06-06 Kabushiki Kaisha Toshiba Steam turbine

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3864056A (en) * 1973-07-27 1975-02-04 Westinghouse Electric Corp Cooled turbine blade ring assembly
GB1524956A (en) * 1975-10-30 1978-09-13 Rolls Royce Gas tubine engine
US5253976A (en) * 1991-11-19 1993-10-19 General Electric Company Integrated steam and air cooling for combined cycle gas turbines
EP1028230A1 (fr) * 1999-02-09 2000-08-16 ABB Alstom Power (Schweiz) AG Pièce refroidie de turbine à gaz avec refroidissement ajustable
US20040219009A1 (en) * 2003-03-06 2004-11-04 Snecma Moteurs Turbomachine with cooled ring segments
EP1526251A1 (fr) * 2003-10-22 2005-04-27 General Electric Company Configuration de refroidissement pour une aube de turbine
US20060140754A1 (en) * 2004-12-27 2006-06-29 Mitsubishi Heavy Industries, Ltd. Gas turbine
EP1923538A2 (fr) * 2006-11-15 2008-05-21 General Electric Company Turbine avec réglage du jeu des extrémités des aubes par transpiration

Also Published As

Publication number Publication date
JP5281166B2 (ja) 2013-09-04
US20110255956A1 (en) 2011-10-20
EP2347100B1 (fr) 2012-10-17
CN102197195B (zh) 2014-03-26
EP2180148A1 (fr) 2010-04-28
CN102197195A (zh) 2011-09-21
JP2012506964A (ja) 2012-03-22
EP2347100A1 (fr) 2011-07-27

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