EP2187002A1 - Agencement d'anneau d'aubes statoriques de turbine à gaz et turbine à gaz - Google Patents

Agencement d'anneau d'aubes statoriques de turbine à gaz et turbine à gaz Download PDF

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Publication number
EP2187002A1
EP2187002A1 EP08019756A EP08019756A EP2187002A1 EP 2187002 A1 EP2187002 A1 EP 2187002A1 EP 08019756 A EP08019756 A EP 08019756A EP 08019756 A EP08019756 A EP 08019756A EP 2187002 A1 EP2187002 A1 EP 2187002A1
Authority
EP
European Patent Office
Prior art keywords
platforms
carrier ring
platform
neighbouring
nozzle arrangement
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Withdrawn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
EP08019756A
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German (de)
English (en)
Inventor
Stephen Batt
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 EP08019756A priority Critical patent/EP2187002A1/fr
Priority to PCT/EP2009/061438 priority patent/WO2010054870A1/fr
Publication of EP2187002A1 publication Critical patent/EP2187002A1/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Withdrawn legal-status Critical Current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
    • F01D11/005—Sealing means between non relatively rotating elements

Definitions

  • the present invention relates to a gas turbine nozzle arrangement comprising an outer carrier ring, an inner carrier ring and nozzle segments each having an outer platform and inner platform and at least one guide vane extending between the outer platform and the inner platform, where the outer platforms each are connected to the outer carrier ring and the inner platforms each are connected to the inner carrier ring.
  • the invention relates to a gas turbine including at least one such nozzle arrangement.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • the turbine blade assembly usually comprises a number of rings of turbine blades between which nozzle arrangements comprising a number of guide vanes are located.
  • a nozzle arrangement typically comprises an outer carrier ring or support ring, an inner carrier ring or support ring, and a number of nozzle segments each comprising a radial outer platform, a radial inner platform and at least one vane extending from the radial outer platform to the radial inner platform.
  • the nozzle arrangement forms an annular flow path for hot and corrosive combustion gases from the combustor.
  • Combustors often operate at high temperatures that may exceed 1350°C.
  • Typical turbine combustor configurations expose turbine vane and blade arrangements to these high temperatures.
  • turbine vanes and blades must be made of materials capable of withstanding such high temperatures.
  • turbine vanes and blades often contain cooling systems for prolonging the lifetime of the vanes and the blades and for reducing the likelihood of failure as a result of excessive temperatures.
  • the platforms are cooled with compressor air.
  • the pressure of the compressor air used for cooling the platforms is higher than the pressure of the combustion gases flowing downstream of the nozzle arrangement.
  • the cooling air used for cooling the platforms, in particular their downstream ends will be discharged into the flow part of the hot combustion gases.
  • the flow of air into the flow path needs to be restricted to a minimum in order to preserve overall turbine efficiency.
  • seals are provided between the radial outer platform of the nozzle segments and the outer carrier ring as well as between the radial inner platform of the nozzle segments and the inner carrier ring.
  • Examples of such seals are disclosed in US 2008/0101927 A1 , US 6,641,144 , US 6,572,331 , US 6,637,753 , US 6,637,751 , US 2005/0244267 A1 , EP 1 323 890 B1 , EP 1 323 896 B1 , EP 1 323 898 B1 , US 6,752,331 , and US 2003/012398 A1 .
  • An inventive gas turbine nozzle arrangement has an axial direction defining a flow direction of hot combustion gas there through and a radial direction.
  • the nozzle arrangement comprises an outer carrier ring, an inner carrier ring, and nozzle segments.
  • Each nozzle segment has an outer platform forming an outer wall segment of a flow channel for the hot combustion gas, an inner platform forming an inner wall segment of a flow channel for the hot combustion gas, and at least one guide vane extending between the outer platform and the inner platform.
  • the outer platforms are each connected to the outer carrier ring and the inner platforms are each connected to the inner carrier ring.
  • the outer carrier ring has an axially facing outer carrier ring surface and/or the inner carrier ring has an axially facing inner carrier ring surface.
  • the outer platform has an axially facing outer platform surface located in axial opposition and at a (small) distance to the axially facing outer carrier ring surface and/or the inner platform has an axially facing inner platform surface located in axial opposition and at a (small) distance to the axially facing inner carrier ring surface.
  • An outer notch extends in a direction towards the centre of the nozzle arrangement at an inclination angle with respect to the radial direction from the outer platform surface into the body of the outer platform.
  • an inner notch extends in a direction towards the centre of the nozzle arrangement at an inclination angle with respect to the radial direction from the outer platform surface into the body of the inner platform.
  • An outer seal strip is inserted into the outer notch, the outer seal strip being dimensioned such that it projects over the outer platform surface and contacts the outer carrier ring surface, and/or an inner seal strip is inserted into the inner notch, the inner seal strip being dimensioned such that it projects over the inner platform surface and contacts the inner carrier ring surface.
  • a number of outer seal strips may be present forming together a conical outer seal strip ring where each outer seal strip is inserted into the outer notches of two or more neighbouring outer platforms.
  • a number of inner seal strips may be present forming together a conical inner seal strip ring where each inner seal strip is inserted into the inner notches of two or more neighbouring inner platforms. Providing a number of seal strips forming together a conical seal strip ring simplifies the assembling of a nozzle segment with a conical seal strip ring.
  • the inventive design of the nozzle segment allows for sealing the leak path between the nozzle segment and the carrier ring with very little complexity and cost. It is suitable for sealing the leak path between the nozzle segment and the outer carrier ring as well as for sealing the leak path between the nozzle segment and the inner carrier ring.
  • the outer platforms and the inner platforms each have peripheral surfaces facing substantially in circumferential direction of the gas turbine nozzle arrangement and being in opposition to a peripheral surface of circumferentially neighbouring ones of the outer platforms and the inner platforms, respectively.
  • a gap is formed between each two of such neighbouring peripheral surfaces.
  • Seals extend in substantially axial direction between the peripheral surfaces of neighbouring outer platforms and/or between the peripheral surfaces of neighbouring inner platforms.
  • the outer seal strip comprises seal strip sections which extend into the gaps between two neighbouring peripheral surfaces of the outer platforms and which are dimensioned such that they contact the axially extending seals between the peripheral surfaces of neighbouring outer platforms and extend over the gap in circumferential direction.
  • the inner seal strip comprises seal strip sections which extend into the gaps between to neighbouring peripheral surfaces of the inner platforms and which are dimensioned such that they contact the axially extending seals between the peripheral surfaces of neighbouring inner platforms and extend over the gap in circumferential direction.
  • the seal strip has a dual function, namely to seal the potential leak between the nozzle and the carrier ring and to seal the leak path between adjacent nozzles segments without a need for additional parts.
  • the outer notch may be elongated in the peripheral surfaces of neighbouring outer platforms so as to reach to the axially extending seals between the respective peripheral surfaces and/or the inner notch is elongated in the peripheral surfaces of neighbouring inner platforms so as to reach the axially extending seals between the respective peripheral surfaces.
  • These elongated notch sections are useful for guiding and holding the mentioned seal strip sections.
  • An inventive gas turbine comprises at least one inventive nozzle arrangement.
  • the same advantages described with respect to the inventive nozzle arrangement are achieved by the inventive gas turbine, in particular sealing the potential air leak path between the nozzle and the carrier ring with little complexity and cost.
  • Figure 1 shows, in a highly schematic view, a gas turbine engine 1 comprising a compressor section 3, a combustor section 5 and a turbine section 7.
  • a rotor 9 extends through all sections and carries, in the compressor section 3, rings of compressor blades 11 and, in the turbine section 7, rings of turbine blades 13. Between neighbouring rings of compressor blades 11 and between neighbouring rings of turbine blades 13, rings of compressor vanes 15 and turbine vanes 17, respectively, extend from a housing 19 of the gas turbine engine 1 radially inwards towards the rotor 9.
  • air is taken in through an air inlet 21 of the compressor section 3.
  • the air is compressed and led towards the combustor section 5 by the rotating compressor blades 11.
  • the air is mixed with a gaseous or liquid fuel and the mixture is burnt.
  • the hot and pressurised combustion gas resulting from burning the fuel/air mixture is fed to the turbine section 7.
  • the hot pressurised gas transfers momentum to the turbine blades 13 while expanding and cooling, thereby imparting a rotation movement to the rotor 9 that drives the compressor and a consumer (not shown), e.g. a generator for producing electrical power or an industrial machine.
  • the rings of turbine vanes 17 function as nozzles for guiding the hot and pressurised combustion gas so as to optimise the momentum transfer to the turbine blades 13.
  • the expanded and cooled combustion gas leaves the turbine section 7 through an exhaust 23.
  • the entrance of the turbine section 7 is shown in more detail in Figure 2 .
  • the figure shows the first ring of turbine blades 13 and a first ring of turbine vanes 17.
  • the turbine vanes 17 extend between radial outer platforms 25 and radial inner platforms 27 that form walls of a flow path for the hot pressurised combustion gas together with neighbouring turbine components 31, 33 and with platforms of the turbine blades 13.
  • Also shown in the figure is the axial direction A and the radial direction R of the rings of turbine vanes and blades. Combustion gas flows through the flow path in the direction indicated in Figure 2 by the arrow 35.
  • the turbine vanes which form nozzle segments together with the outer and inner platform between which they extend, are held in place by an outer carrier ring 37 and an inner carrier ring 39 to which the outer platforms and the inner platforms, respectively, are connected.
  • the outer carrier ring 37, the inner carrier ring 39 and the nozzle segments together form a nozzle arrangement of the turbine.
  • each single guide vane 17 of the present embodiment forms a nozzle segment together with the outer platform 25 and the inner platform 27
  • the outer platform and an inner platform could extend over a larger ring segment than in the depicted embodiment so that they could have a number of vanes, e.g., two or three vanes, extending between them.
  • platforms extending over a smaller ring segment and having only one vane extending between them are advantageous as thermal expansion during gas turbine operation leads to less internal stress than with platforms extending over a larger ring segment.
  • Figures 3 and 4 show the nozzle arrangement depicted in Figure 2 in more detail. While Figure 3 shows a section of the nozzle arrangement in a perspective view Figure 4 shows the same section in a sectional view, including sections of the radial outer platform 25 and the outer carrier ring 37.
  • the radial outer platform 25 comprises a rail 41 with an axially facing platform surface 43. Furthermore, the outer platform 25 comprises peripheral surfaces 45 facing substantially in circumferential direction of the nozzle arrangement. In the nozzle arrangement, gaps are present between the peripheral surfaces 45 of neighbouring outer platforms. Furthermore, the peripheral surfaces 45 are equipped with grooves 47 for receiving seal strips extending between the neighbouring platforms for sealing the gap there between.
  • the carrier ring 37 comprises an axially facing surface 49 located in axial opposition to the axially facing surface 43 of the outer platforms rail 41.
  • a bulge 44 may be part of the outer carrier rings axially facing surface 49 or, as in the present embodiment, may be part of the axially facing surface 43 of the rail. The bulge 44 provides for a defined distance between both opposed surfaces 43 and 49.
  • the outer platform 25 is provided with a notch 51 in the axially facing surface 43 of the rail 41.
  • This notch extends circumferentially through the axially facing surface 43.
  • the notch extends along a direction that is inclined towards the flow path through the nozzle arrangement, i.e. towards the centre of the nozzle arrangement, with respect to the radial direction of the nozzle arrangement into the body of the outer platform 25. Its depth is elongated in the peripheral surfaces 45 of the platform 25 so that it reaches the grooves 47 in the peripheral surfaces and itself forms inclined grooves in these surfaces.
  • the nozzle arrangement further comprises seal strips 53 fitted into the notch 51.
  • the dimension of a seal strip 53 is chosen such that it projects over the axially surface 43 of the rail 41 when it is inserted into the notch 51 by such an amount that it contacts the axially facing surface 49 of the outer carrier ring 37 and, hence, seals the gap between the opposing surfaces of the outer carrier ring 37 and the rail 41.
  • a number of such seal strips 53 forms a conical seal strip ring of the nozzle arrangement.
  • a seal strip 53 which is schematically shown in Figure 5 , further comprises elongated seal strip sections 55 which extend into the gaps between neighbouring outer platforms 25 and which are dimensioned such that they contact the seals extending between the peripheral surfaces 45 of neighbouring outer platforms 25 with their axial ends 57.
  • the seal strip 53 comprises extended sections distributed over its circumferential direction, the lengths of which are chosen such as to reach the seal between neighbouring platforms when inserted into the notch and the width of which correspond to the sum of the gap width between neighbouring platforms and the depth of the elongated sections of the grooves in opposing peripheral surfaces 45.
  • seal strip of the embodiment shown in figure 5 extends over more than two platforms 25 of a nozzle segment (for example three or four platforms) it may be advantageous in view of assembling a nozzle segment with a conical seal ring if the seal strip ring is composed of shorter seal strips which only extend over two circumferentially neighbouring platforms.
  • An embodiment with several shorter seal strips extending only over two neighbouring platforms is schematically shown for the inner platforms 27 of a nozzle segment in figure 6 .
  • Figure 6 shows a number of inner platforms 27 which form an inner platform ring.
  • a seal strip ring composed of a number of seal strips 153A, 153B, 153C is inserted into notches of the inner platforms 27.
  • Each seal strip 153A, 153B, 153C extends over half of a platform's 27 circumferential extension.
  • Each seal strip 153A, 153B, 153C looks roughly like the form of the letter T.
  • a not assembled seal strip 53 is shown in figure 6 on the right hand side next to the assembled ones to show the proportions.
  • Small gaps 131 may be present between the seal strips 153A, 153B, 153C in circumferential direction when inserted into the notches of the inner platform 27 to allow for thermal expansion. However, if thermal expansion is negligible or can be accounted for by other means such gaps may be omitted.
  • Each seal strip 153A, 153B, 153C comprises an elongated seal strip section 155A, 155B, 155C which is located at the seals strip's centre and extends into a gap between two neighbouring inner platforms 27.
  • the seal strip section 155A, 155B, 155C is dimensioned such that it contacts seals 159 extending between the peripheral surfaces of neighbouring inner platforms 27 with their axial ends 157.
  • the seals 159 and the elongated seal strip sections 155A, 155B, 155C may both end at the same spot indicated by reference sign 157 (not shown) or the seals 159 may be of longer length and for example end at position 162, whereas the elongated seal strip section 155A, 155B, 155C end at position 157.
  • Dashed lines 160 indicate grooves (reference sign 47 in figure 4 ) that hold the contacts seals 159 between two neighbouring inner platforms 27 and hold the elongated seal strip section 155A, 155B, 155C.
  • the further dashed lines 161 indicate notches (reference signs 51 in the previous figures) to hold the circumferential part of the seal strips 153A, 153B, 153C.
EP08019756A 2008-11-12 2008-11-12 Agencement d'anneau d'aubes statoriques de turbine à gaz et turbine à gaz Withdrawn EP2187002A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP08019756A EP2187002A1 (fr) 2008-11-12 2008-11-12 Agencement d'anneau d'aubes statoriques de turbine à gaz et turbine à gaz
PCT/EP2009/061438 WO2010054870A1 (fr) 2008-11-12 2009-09-04 Configuration de tuyère pour turbine à gaz, et turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP08019756A EP2187002A1 (fr) 2008-11-12 2008-11-12 Agencement d'anneau d'aubes statoriques de turbine à gaz et turbine à gaz

Publications (1)

Publication Number Publication Date
EP2187002A1 true EP2187002A1 (fr) 2010-05-19

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EP08019756A Withdrawn EP2187002A1 (fr) 2008-11-12 2008-11-12 Agencement d'anneau d'aubes statoriques de turbine à gaz et turbine à gaz

Country Status (2)

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EP (1) EP2187002A1 (fr)
WO (1) WO2010054870A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN108980346A (zh) * 2017-05-30 2018-12-11 森萨塔科技公司 垫圈固定方法和系统

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102019108267A1 (de) * 2019-03-29 2020-10-01 Rolls-Royce Deutschland Ltd & Co Kg Vorrichtung zur Befestigung von Dichtplatten zwischen Bauteilen eines Gasturbinentriebwerks
DE102019113530A1 (de) * 2019-05-21 2020-11-26 Rolls-Royce Deutschland Ltd & Co Kg Vorrichtung zur Abdichtung eines Spalts zwischen zwei Bauteilen einer Turbine eines Gasturbinentriebwerks

Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2000012870A1 (fr) * 1998-09-02 2000-03-09 General Electric Company Joint d'etancheite annulaire en forme de c
US6164656A (en) * 1999-01-29 2000-12-26 General Electric Company Turbine nozzle interface seal and methods
US20030012398A1 (en) 2001-07-11 2003-01-16 Sunshine Jessica Miriam Method for selecting representative endmember components from spectral data
US6572331B1 (en) 2001-12-28 2003-06-03 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
EP1323896A2 (fr) * 2001-12-28 2003-07-02 General Electric Company Joint statorique pour turbine à gaz
US6637753B2 (en) 2001-12-28 2003-10-28 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
US6637751B2 (en) 2001-12-28 2003-10-28 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
US6641144B2 (en) 2001-12-28 2003-11-04 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
US6752331B2 (en) 2002-04-17 2004-06-22 Sk & Y Agricultural Equipments Co., Ltd. Air-pressure sprayer structure
US20050244267A1 (en) 2004-04-29 2005-11-03 General Electric Company System for sealing an inner retainer segment and support ring in a gas turbine and methods therefor
EP1323890B1 (fr) 2001-12-28 2006-09-27 General Electric Company Turbine comprenant un joint auxiliaire pour l' étanchéité d'éléments statoriques
US20080101927A1 (en) 2006-10-25 2008-05-01 Siemens Power Generation, Inc. Turbine vane ID support

Family Cites Families (1)

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Publication number Priority date Publication date Assignee Title
FR2829796B1 (fr) * 2001-09-20 2003-12-12 Snecma Moteurs Dispositif de maintien des joints de plates-formes de secteurs de distributeur de turbomachine a lamelles d'etancheite

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2000012870A1 (fr) * 1998-09-02 2000-03-09 General Electric Company Joint d'etancheite annulaire en forme de c
US6164656A (en) * 1999-01-29 2000-12-26 General Electric Company Turbine nozzle interface seal and methods
US20030012398A1 (en) 2001-07-11 2003-01-16 Sunshine Jessica Miriam Method for selecting representative endmember components from spectral data
US6637751B2 (en) 2001-12-28 2003-10-28 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
EP1323896A2 (fr) * 2001-12-28 2003-07-02 General Electric Company Joint statorique pour turbine à gaz
US6637753B2 (en) 2001-12-28 2003-10-28 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
US6572331B1 (en) 2001-12-28 2003-06-03 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
US6641144B2 (en) 2001-12-28 2003-11-04 General Electric Company Supplemental seal for the chordal hinge seals in a gas turbine
EP1323898B1 (fr) 2001-12-28 2006-06-07 General Electric Company Joint statorique pour turbine à gaz
EP1323896B1 (fr) 2001-12-28 2006-08-16 General Electric Company Turbine ayant un joint statorique
EP1323890B1 (fr) 2001-12-28 2006-09-27 General Electric Company Turbine comprenant un joint auxiliaire pour l' étanchéité d'éléments statoriques
US6752331B2 (en) 2002-04-17 2004-06-22 Sk & Y Agricultural Equipments Co., Ltd. Air-pressure sprayer structure
US20050244267A1 (en) 2004-04-29 2005-11-03 General Electric Company System for sealing an inner retainer segment and support ring in a gas turbine and methods therefor
US20080101927A1 (en) 2006-10-25 2008-05-01 Siemens Power Generation, Inc. Turbine vane ID support

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN108980346A (zh) * 2017-05-30 2018-12-11 森萨塔科技公司 垫圈固定方法和系统

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