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
• • 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
  • 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 noz- zle 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 0 C.
• Typical turbine combustor configurations expose turbine vane and blade arrangements to these high temperatures.
• turbine vanes and blades must be made of materi- als 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 .
• 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 radial inner platform, in particular its downstream end will be discharged into the flow part of the hot combustion gases.
• the flow of cooling air into the flow path needs to be restricted to a minimum in order to preserve overall turbine efficiency.
• 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 Al, US 6,641,144, US 6,572,331, US 6, .637, 753, US 6,637,751 und US 2005/0244267 Al.
• 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 sup- port, 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 out- wards 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 seg- ments 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, re- spectively, 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 tur- bine 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 pressur- ised 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 em- bodiment 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 em- bodiment 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 in- ner 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 ra- dially 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 be- tween 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.
• the cooling air passing through the through holes 67 in the seal strip 41 forms impingement jets impinging onto the radial inner 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.
• leakage of cooling air into the flow path of the hot combustion gases can be kept small.
• 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)

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Citations (4)

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• 2008
  • 2008-10-23 EP EP08018594A patent/EP2180143A1/de not_active Withdrawn
• 2009
  • 2009-08-27 WO PCT/EP2009/061050 patent/WO2010046167A1/en active Application Filing

Patent Citations (4)

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