EP2180143A1 - Gas turbine nozzle arrangement and gas turbine - Google Patents
Gas turbine nozzle arrangement and gas turbine Download PDFInfo
- Publication number
- EP2180143A1 EP2180143A1 EP08018594A EP08018594A EP2180143A1 EP 2180143 A1 EP2180143 A1 EP 2180143A1 EP 08018594 A EP08018594 A EP 08018594A EP 08018594 A EP08018594 A EP 08018594A EP 2180143 A1 EP2180143 A1 EP 2180143A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- ring section
- gas turbine
- platforms
- rails
- carrier ring
- 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- Withdrawn
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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
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
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- 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
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- 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
- F05D2240/80—Platforms for stationary or moving blades
- F05D2240/81—Cooled platforms
-
- 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 present invention relates to a gas turbine nozzle arrangement comprising an outer support, a 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 support and the inner platforms each are connected to the 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 support, 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 reducing the likelihood of failure as a result of excessive temperatures.
- An inventive gas turbine nozzle arrangement has an axial direction defining a flow direction of hot combustion gas there through and a radial direction. It comprises an outer support, a carrier ring, and nozzle segments.
- the carrier ring comprises a carrier ring section which extends radially outwards and has a radially outer surface, i.e. a surface normal of which shows radially outwards.
- the nozzle segments each have an outer platform, an inner platform, and at least one guide vane extending between the outer platform and the inner platform.
- the outer platforms of the nozzle segments each have a radial inner surface forming an outer flow channel wall for the hot combustion gas
- the inner platforms of the nozzle segments each have a radially outer surface forming an inner flow channel wall for the hot combustion gas.
- the inner platforms comprise a downstream end with respect to the flow direction of the hot combustion gas, a radially inner surface (the surface normal of which showing radially inwards) and a rail extending radially from the radial inner surface.
- the outer platforms are each connected to the outer support while the inner platforms are each connected to the carrier ring by means of the rail and the ring section such that the rail overlaps the ring section, in particular such that the rails are located upstream of the carrier ring section.
- the overlap may specifically be arranged that way, that the rail extends radially inwards and the ring section radially outwards and that a part of the rail and a part of the ring section will be adjacent to each other.
- At least one flow channel for cooling fluid for example compressor air, is formed between the rails and the ring section.
- at least one seal strip is present between the radially outer surface of the carrier ring section and the inner surface of the inner platforms.
- the seal strip comprises through holes for allowing cooling fluid to flow to the sealing strip.
- the air leak between nozzle and carrier ring is kept below a certain amount. Moreover, the flow of air which is allowed to pass the seal strip is controlled by way of the holes in the seal strip. In other words, the inventive nozzle arrangement allows for a high degree in controlling the air which is allowed to pass the seal strip, for cooling the inner platforms, in particular the platforms downstream ends.
- the through holes can be used for forming cooling fluid jets so as to provide for impingement cooling of the inner surface of the platform, in particular close to its downstream edge.
- the ring section and the rails may abut on each other, each comprising an abutting surface.
- at least one flow channel is formed by grooves provided in the inner ring segment's abutting surface and/or the rails' abutting surfaces.
- the carrier ring section may comprise a number of blind holes, and the rails each may comprise at least one through hole.
- the inner platforms then can be connected to the carrier ring by means of bolts extending through the through holes of the rails into the blind holes of the carrier ring section.
- An inventive gas turbine comprises at least one inventive gas turbine nozzle arrangement.
- the inventive nozzle arrangement allows for highly controlling the leakage between nozzle segments and the carrier ring and to provide for effective impingement cooling of the inner platforms without the need of additional parts.
- Figure 1 shows a gas turbine engine in a highly schematic view.
- Figure 2 shows the turbine entry of a gas turbine engine.
- Figure 3 shows a section of the inventive nozzle arrangement in a perspective view.
- Figure 4 shows the section of figure 3 in a sectional view.
- 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 - the part closest to the combustor section 5 - 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 17, which form nozzle segments together with the outer and inner platform between which they extend, are held in place by an outer support 37 and an inner support 39, the latter called carrier ring in the following, to which the outer platforms and the inner platforms, respectively, are connected.
- each single guide vane of the present embodiment forms a nozzle segment together with the outer platform 25 and the inner platform 27
- the outer platform 25 and an inner platform 27 could extend over a larger ring segment than in the depicted embodiment and 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 prospective view figure 4 shows the same section in a sectional view. Both views show sections of the radial inner platform 27 and the carrier ring 39. Also visible is a seal strip 41 located between a radial inner surface 43 of the platform's 27 downstream end and a radial outer surface 47 of a radially extending ring section 45 of the carrier ring 39.
- the inner platform 27 is fixed to the carrier ring 39 by means of a rail 49 that extends radially inwards from the radial inner surface 43 of the inner platform 27 and abuts on the upstream side of the ring section 45 of the carrier ring 39.
- the rail 49 and the ring section 45 have plain abutting surfaces 51, 53 with one or more channels 55 present in at least one of these abutting surfaces.
- flow channels 55 are present in the rail's 49 abutting surface 53.
- Bolts 57 extend through through holes 59 in the rail 49 into blind holes 61 in the ring section 45 and fix the rail 49 to the ring section 45.
- the seal strip 41 is held in place between the radial inner surface 43 of the platform 27 and the radial outer surface 47 of the ring section 45 by a projection 63 projecting radially outwards from ring section's 45 radial outer surface 47.
- the projection 63 only projects over the surface 47 radially outwards by an amount which leaves a gap 65 between the projection 63 and the radial inner surface 43 of the platform's 27 downstream section when the nozzle segment is fixed to the carrier ring 39.
- a cavity 68 with a downstream flow exit to the flow path of the hot combustion gas is present between the radial outer surface 47 of the carrier ring's ring section 45 and the radial inner surface 43 of the platform's downstream sections which can accommodate the seal strip 41.
- the seal strip 41 comprises through holes 67 - as openings through the seal strip 41 - which form, together with the flow channels 55, the cavity 68 and the gaps 65, a flow path for allowing cooling air to flow from a space 69 formed between the platform 27 and the carrier ring 39 towards and along the radial inner surface 43 of the platform's downstream ends 28, hence cooling the downstream ends 28.
- through holes 67 in form of bores are used in the present embodiment other shapes of through holes, like long holes, or slots in the seal strip 41, could be used as well.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
an outer support (37), a carrier ring (39) comprising a carrier ring section (45) extending radially outwards and having a radially outer surface (47), and nozzle segments each having an outer platform (25), an inner platform (27) and at least one guide vane (17) extending between the outer platform (25) and the inner platform (27). The outer platforms (25) of the nozzle segments form an outer flow channel wall for the hot combustion gas. The inner platforms (27) of the nozzle segments form an inner flow channel wall for the hot combustion gas and each comprise a downstream end (28) with respect to the flow direction, a radially inner surface (43) and a rail (49) extending radially inwards from the radially inner surface (43). While the outer platforms (25) each are connected to the outer support (37) the inner platforms (27) each are connected to the carrier ring (39) by means of the rails (49) and the ring section (45) such that the rails (49) overlap the carrier ring section (45). At least one flow channel (55) for a cooling fluid is formed between the rails (49) and the ring section (45). In addition, at least one seal strip (41) is present between the radially outer surface (47) of the carrier ring section (45) and inner surface (43) of the inner platforms (27) and comprises openings (67) for allowing cooling fluid to flow through the seal strip (41)
Description
- The present invention relates to a gas turbine nozzle arrangement comprising an outer support, a 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 support and the inner platforms each are connected to the carrier ring. In addition, the invention relates to a gas turbine including at least one such nozzle arrangement.
- Typically, gas turbine engines 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 support, 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. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for prolonging the lifetime of the vanes and the blades and reducing the likelihood of failure as a result of excessive temperatures.
- In order to prevent the platforms of the nozzle segments, which form the walls of the flow path for the hot and corrosive combustion gases, from damage due to the hot combustion gases the platforms are cooled with compressor air. However, 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. Moreover, the cooling air used for cooling the radial inner platform, in particular its downstream end, will be discharged into the flow part of the hot combustion gases. Hence, the flow of cooling air into the flow path needs to be restricted to a minimum in order to preserve overall turbine efficiency. Therefore, seals are provided at the radial inner platform of the nozzle segments and the carrier ring in order to restrict the flow of compressor air into the flow path of the hot combustion gas. 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 undUS 2005/0244267 A1 . - With the respect to the mentioned prior art it is an objective of the present invention to provide an advantageous gas turbine nozzle arrangement and an advantageous gas turbine.
- These objectives are solved by gas turbine nozzle arrangements as claimed in claim 1 and by a gas turbine as claimed in
claim 7. The depending claims contain further the developments of the invention. - An inventive gas turbine nozzle arrangement has an axial direction defining a flow direction of hot combustion gas there through and a radial direction. It comprises an outer support, a carrier ring, and nozzle segments.
- The carrier ring comprises a carrier ring section which extends radially outwards and has a radially outer surface, i.e. a surface normal of which shows radially outwards.
- The nozzle segments each have an outer platform, an inner platform, and at least one guide vane extending between the outer platform and the inner platform. The outer platforms of the nozzle segments each have a radial inner surface forming an outer flow channel wall for the hot combustion gas, while the inner platforms of the nozzle segments each have a radially outer surface forming an inner flow channel wall for the hot combustion gas. Moreover, the inner platforms comprise a downstream end with respect to the flow direction of the hot combustion gas, a radially inner surface (the surface normal of which showing radially inwards) and a rail extending radially from the radial inner surface.
- The outer platforms are each connected to the outer support while the inner platforms are each connected to the carrier ring by means of the rail and the ring section such that the rail overlaps the ring section, in particular such that the rails are located upstream of the carrier ring section. The overlap may specifically be arranged that way, that the rail extends radially inwards and the ring section radially outwards and that a part of the rail and a part of the ring section will be adjacent to each other. At least one flow channel for cooling fluid, for example compressor air, is formed between the rails and the ring section. Moreover, at least one seal strip is present between the radially outer surface of the carrier ring section and the inner surface of the inner platforms. The seal strip comprises through holes for allowing cooling fluid to flow to the sealing strip.
- In the inventive design of the nozzle arrangement, the air leak between nozzle and carrier ring is kept below a certain amount. Moreover, the flow of air which is allowed to pass the seal strip is controlled by way of the holes in the seal strip. In other words, the inventive nozzle arrangement allows for a high degree in controlling the air which is allowed to pass the seal strip, for cooling the inner platforms, in particular the platforms downstream ends.
- When the seal strip is inclined with respect to the radial direction of the carrier ring, in particular, if the inclination angle of the seal strips inclination is larger than 60 degree, the through holes can be used for forming cooling fluid jets so as to provide for impingement cooling of the inner surface of the platform, in particular close to its downstream edge.
- For connecting the inner platforms to the carrier ring the ring section and the rails may abut on each other, each comprising an abutting surface. In this case at least one flow channel is formed by grooves provided in the inner ring segment's abutting surface and/or the rails' abutting surfaces.
- For fixing the rails to the carrier ring the carrier ring section may comprise a number of blind holes, and the rails each may comprise at least one through hole. The inner platforms then can be connected to the carrier ring by means of bolts extending through the through holes of the rails into the blind holes of the carrier ring section.
- An inventive gas turbine comprises at least one inventive gas turbine nozzle arrangement. The inventive nozzle arrangement allows for highly controlling the leakage between nozzle segments and the carrier ring and to provide for effective impingement cooling of the inner platforms without the need of additional parts.
-
Figure 1 shows a gas turbine engine in a highly schematic view. -
Figure 2 shows the turbine entry of a gas turbine engine. -
Figure 3 shows a section of the inventive nozzle arrangement in a perspective view. -
Figure 4 shows the section offigure 3 in a sectional view. -
Figure 1 shows, in a highly schematic view, a gas turbine engine 1 comprising acompressor section 3, a combustor section 5 and aturbine section 7. A rotor 9 extends through all sections and carries, in thecompressor section 3, rings of compressor blades 11 and, in theturbine section 7, rings ofturbine blades 13. Between neighbouring rings of compressor blades 11 and between neighbouring rings ofturbine blades 13, rings ofcompressor vanes 15 andturbine vanes 17, respectively, extend from ahousing 19 of the gas turbine engine 1 radially inwards towards the rotor 9. - In operation of the gas turbine engine 1 air is taken in through an
air inlet 21 of thecompressor section 3. The air is compressed and led towards the combustor section 5 by the rotating compressor blades 11. In the combustor section 5 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 theturbine section 7. On its way through theturbine section 7 the hot pressurised gas transfers momentum to theturbine 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 theturbine blades 13. Finally, the expanded and cooled combustion gas leaves theturbine section 7 through anexhaust 23. - The entrance of the turbine section 7 - the part closest to the combustor section 5 - is shown in more detail in
Figure 2 . The figure shows the first ring ofturbine blades 13 and a first ring of turbine vanes 17. The turbine vanes 17 extend between radialouter platforms 25 and radialinner platforms 27 that form walls of a flow path for the hot pressurised combustion gas together with neighbouringturbine components 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 inFigure 2 by thearrow 35. The turbine vanes 17, which form nozzle segments together with the outer and inner platform between which they extend, are held in place by anouter support 37 and aninner support 39, the latter called carrier ring in the following, to which the outer platforms and the inner platforms, respectively, are connected. Theouter support 37, thecarrier ring 39 and the nozzle segments together form a nozzle arrangement of the turbine. - Note, that although each single guide vane of the present embodiment forms a nozzle segment together with the
outer platform 25 and theinner platform 27 other forms of nozzle segments may be possible. In an exemplary alternative nozzle segment, theouter platform 25 and aninner platform 27 could extend over a larger ring segment than in the depicted embodiment and could have a number of vanes, e.g., two or three vanes, extending between them. However, 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 and4 show the nozzle arrangement depicted infigure 2 in more detail. Whilefigure 3 shows a section of the nozzle arrangement in a prospective viewfigure 4 shows the same section in a sectional view. Both views show sections of the radialinner platform 27 and thecarrier ring 39. Also visible is aseal strip 41 located between a radialinner surface 43 of the platform's 27 downstream end and a radialouter surface 47 of a radially extendingring section 45 of thecarrier ring 39. Theinner platform 27 is fixed to thecarrier ring 39 by means of arail 49 that extends radially inwards from the radialinner surface 43 of theinner platform 27 and abuts on the upstream side of thering section 45 of thecarrier ring 39. Therail 49 and thering section 45 have plain abuttingsurfaces more channels 55 present in at least one of these abutting surfaces. In the present embodiment,flow channels 55 are present in the rail's 49 abuttingsurface 53.Bolts 57 extend through throughholes 59 in therail 49 intoblind holes 61 in thering section 45 and fix therail 49 to thering section 45. - The
seal strip 41 is held in place between the radialinner surface 43 of theplatform 27 and the radialouter surface 47 of thering section 45 by aprojection 63 projecting radially outwards from ring section's 45 radialouter surface 47. However, theprojection 63 only projects over thesurface 47 radially outwards by an amount which leaves agap 65 between theprojection 63 and the radialinner surface 43 of the platform's 27 downstream section when the nozzle segment is fixed to thecarrier ring 39. Hence, acavity 68 with a downstream flow exit to the flow path of the hot combustion gas is present between the radialouter surface 47 of the carrier ring'sring section 45 and the radialinner surface 43 of the platform's downstream sections which can accommodate theseal strip 41. - The
seal strip 41 comprises through holes 67 - as openings through the seal strip 41 - which form, together with theflow channels 55, thecavity 68 and thegaps 65, a flow path for allowing cooling air to flow from aspace 69 formed between theplatform 27 and thecarrier ring 39 towards and along the radialinner surface 43 of the platform's downstream ends 28, hence cooling the downstream ends 28. Due to the inclination of theseal strip 41 with respect to radial R direction of the nozzle arrangement, which is in the present embodiment about 60 degree, the cooling air passing through the throughholes 67 in theseal strip 41 forms impingement jets impinging onto the radialinner surface 43 of the platform's downstream ends 28, which increases the cooling efficiency and hence allows to reduce the amount of cooling air necessary for effectively cooling the downstream ends 28. As a consequence, leakage of cooling air into the flow path of the hot combustion gases can be kept small. - Note that although through
holes 67 in form of bores are used in the present embodiment other shapes of through holes, like long holes, or slots in theseal strip 41, could be used as well.
Claims (7)
- A gas turbine nozzle arrangement, having an axial direction (A) defining a flow direction of hot combustion gas there through and a radial direction (R), the nozzle arrangement comprising:- an outer support (37),- a carrier ring (39) comprising a carrier ring section (45) extending radially outwards and having a radially outer surface (47), and- nozzle segments each having an outer platform (25), an inner platform (27) and at least one guide vane (17) extending between the outer platform (25) and the inner platform (27),- the outer platforms (25) of the nozzle segments forming an outer flow channel wall for the hot combustion gas,- the inner platforms (27) of the nozzle segments forming an inner flow channel wall for the hot combustion gas and each comprising a downstream end (28) with respect to the flow direction, a radially inner surface (43) and a rail (49) extending radially inwards from the radially inner surface (43),- where the outer platforms (25) each are connected to the outer support (37) and the inner platforms (27) each are connected to the carrier ring (39) by means of the rails (49) and the ring section (45) such that the rails (49) overlap the ring section (45),characterised in that- at least one flow channel (55) for a cooling fluid is formed between the rails (49) and the ring section (45) and- at least one seal strip (41) is present between the radially outer surface (47) of the carrier ring section (45) and the radially inner surface (43) of the inner platforms (27), which seal strip (41) comprises openings (67) for allowing cooling fluid to flow through the seal strip (41).
- The gas turbine nozzle arrangement as claimed in claim 1,
characterised in that
the inner platforms (27) each are connected to the carrier ring (39) by means of the rails (49) and the ring section (45) such that the rails (49) overlap the ring section (45) upstream of the ring section (45). - The gas turbine nozzle arrangement as claimed in claim 1 or claim 2,
characterised in that
the seal strip (41) is inclined with respect to the radial direction (R). - The gas turbine nozzle arrangement as claimed in claim 3,
characterised in that the inclination angle of the seal strip's (41) inclination is at least 60 degree. - The gas turbine nozzle arrangement as claimed in any one of the claims 1 to 4,
characterised in that- the ring section (45) and the rails (49) abut on each other and each comprise an abutting surface (51, 53),- the at least one flow channel (55) is formed by grooves formed in the ring segment's abutting surface (51) and/or the rails' abutting surfaces (53). - The gas turbine nozzle arrangement as claimed in any one of the claims 1 to 5,
characterised in that- the carrier ring section (45) comprises a number of blind holes (61),- the rails (49) each comprise at least one through hole (59), and- the inner platforms (27) are connected to the carrier ring (39) by means of bolts (57) extending through the through holes (59) of the rails (49) into the blind holes (61) of the carrier ring section (45). - A gas turbine comprising at least one gas turbine nozzle arrangement as claimed in any one of the claims 1 to 6.
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP08018594A EP2180143A1 (en) | 2008-10-23 | 2008-10-23 | Gas turbine nozzle arrangement and gas turbine |
PCT/EP2009/061050 WO2010046167A1 (en) | 2008-10-23 | 2009-08-27 | Gas turbine nozzle arrangement and gas turbine |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP08018594A EP2180143A1 (en) | 2008-10-23 | 2008-10-23 | Gas turbine nozzle arrangement and gas turbine |
Publications (1)
Publication Number | Publication Date |
---|---|
EP2180143A1 true EP2180143A1 (en) | 2010-04-28 |
Family
ID=40451255
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP08018594A Withdrawn EP2180143A1 (en) | 2008-10-23 | 2008-10-23 | Gas turbine nozzle arrangement and gas turbine |
Country Status (2)
Country | Link |
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EP (1) | EP2180143A1 (en) |
WO (1) | WO2010046167A1 (en) |
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP2415969A1 (en) * | 2010-08-05 | 2012-02-08 | Siemens Aktiengesellschaft | Component of a turbine with leaf seals and method for sealing against leakage between a vane and a carrier element |
EP2660429A1 (en) * | 2012-05-03 | 2013-11-06 | Siemens Aktiengesellschaft | Sealing arrangement for a nozzle guide vane and gas turbine |
US11111794B2 (en) | 2019-02-05 | 2021-09-07 | United Technologies Corporation | Feather seals with leakage metering |
Citations (10)
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---|---|---|---|---|
US4883405A (en) * | 1987-11-13 | 1989-11-28 | The United States Of America As Represented By The Secretary Of The Air Force | Turbine nozzle mounting arrangement |
EP0513956A1 (en) * | 1991-05-13 | 1992-11-19 | General Electric Company | Boltless turbine nozzle/stationary seal mounting |
US6572331B1 (en) | 2001-12-28 | 2003-06-03 | 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 |
US6637753B2 (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 |
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 |
EP1607582A1 (en) * | 2004-06-17 | 2005-12-21 | Snecma Moteurs | Mounting of a gas turbine combustor with integrated turbine inlet guide conduit |
US20060127215A1 (en) * | 2004-12-15 | 2006-06-15 | Pratt & Whitney Canada Corp. | Integrated turbine vane support |
US20080101927A1 (en) | 2006-10-25 | 2008-05-01 | Siemens Power Generation, Inc. | Turbine vane ID support |
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2008
- 2008-10-23 EP EP08018594A patent/EP2180143A1/en not_active Withdrawn
-
2009
- 2009-08-27 WO PCT/EP2009/061050 patent/WO2010046167A1/en active Application Filing
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Cited By (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP2415969A1 (en) * | 2010-08-05 | 2012-02-08 | Siemens Aktiengesellschaft | Component of a turbine with leaf seals and method for sealing against leakage between a vane and a carrier element |
WO2012016790A1 (en) | 2010-08-05 | 2012-02-09 | Siemens Aktiengesellschaft | Component of a turbine with leaf seals and method for sealing against leakage between a vane and a carrier element |
CN103052766A (en) * | 2010-08-05 | 2013-04-17 | 西门子公司 | Component of a turbine with leaf seals and method for sealing against leakage between a vane and a carrier element |
RU2583487C2 (en) * | 2010-08-05 | 2016-05-10 | Сименс Акциенгезелльшафт | Turbine component with plate seals and method of sealing against leak between blade and carrying element |
US9506374B2 (en) | 2010-08-05 | 2016-11-29 | Siemens Aktiengesellschaft | Component of a turbine with leaf seals and method for sealing against leakage between a vane and a carrier element |
EP2660429A1 (en) * | 2012-05-03 | 2013-11-06 | Siemens Aktiengesellschaft | Sealing arrangement for a nozzle guide vane and gas turbine |
WO2013164184A1 (en) * | 2012-05-03 | 2013-11-07 | Siemens Aktiengesellschaft | Sealing arrangement for a nozzle guide vane and gas turbine |
CN104334833A (en) * | 2012-05-03 | 2015-02-04 | 西门子公司 | Sealing arrangement for a nozzle guide vane and gas turbine |
CN104334833B (en) * | 2012-05-03 | 2017-04-05 | 西门子公司 | For nozzle guide vanes and the sealing device of gas turbine |
US9617920B2 (en) | 2012-05-03 | 2017-04-11 | Siemens Aktiengesellschaft | Sealing arrangement for a nozzle guide vane and gas turbine |
US11111794B2 (en) | 2019-02-05 | 2021-09-07 | United Technologies Corporation | Feather seals with leakage metering |
Also Published As
Publication number | Publication date |
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WO2010046167A1 (en) | 2010-04-29 |
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