WO2018128609A1 - Ensemble joint d'étanchéité entre un trajet de gaz chaud et une cavité de disque de rotor - Google Patents

Ensemble joint d'étanchéité entre un trajet de gaz chaud et une cavité de disque de rotor Download PDF

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Publication number
WO2018128609A1
WO2018128609A1 PCT/US2017/012262 US2017012262W WO2018128609A1 WO 2018128609 A1 WO2018128609 A1 WO 2018128609A1 US 2017012262 W US2017012262 W US 2017012262W WO 2018128609 A1 WO2018128609 A1 WO 2018128609A1
Authority
WO
WIPO (PCT)
Prior art keywords
face
sidewall
seal
groove floor
radially
Prior art date
Application number
PCT/US2017/012262
Other languages
English (en)
Inventor
Bharat Sanjay PRABHU
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 PCT/US2017/012262 priority Critical patent/WO2018128609A1/fr
Publication of WO2018128609A1 publication Critical patent/WO2018128609A1/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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/082—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades on the side of the rotor disc
    • 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
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • 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
    • F05D2260/14—Preswirling

Definitions

  • the present invention is directed generally to gas turbine engines, and in particular, to a seal assembly for assisting in limiting leakage between a hot gas path and a rotor disc cavity in the turbine section of a gas turbine engine.
  • a gas turbine engine typically includes a compressor section, a combustor, and a turbine section.
  • the compressor section typically compresses ambient air that enters an inlet.
  • the combustor combines the compressed air with a fuel and ignites the mixture creating combustion products defining a working fluid.
  • the working fluid travels to the turbine section where it is expanded to produce a work output.
  • rows of stationary flow directing members comprising vanes directing the working fluid to rows of rotating flow directing members comprising blades coupled to a rotor. Each pair of a row of vanes and a row of blades forms a stage in the turbine section.
  • the seal gap between rotating blades and stationary vanes may be prone to hot has ingestion from the hot has path into rotor cavities that contain cooling fluids.
  • hot gas ingestion reduces engine performance and efficiency, e.g., by yielding higher disc and blade root temperatures.
  • Ingestion of the working gas from the hot gas path to the disc cavities may also reduce service life and/or cause failure of the components in and around the disc cavities.
  • cooling fluid such as compressor air may be supplied through the seal gap between the rotating blades and the stationary vanes. While this cooler air may reduce some amount of hot gas ingestion, it may also induce aerodynamic mixing losses while mixing with the hot gas. Aerodynamic mixing losses typically occur due to a difference in circumferential/tangential momentum between the cooling air and the hot gas.
  • aspects of the present invention provide a turbine rotating component with a pumping feature, embodied as grooves on the rotating component, which forms part of a seal assembly between a hot gas path and a rotor disc cavity in a gas turbine engine.
  • the seal assembly counteracts hot gas ingestion into the rotor disc cavity from the hot gas path.
  • a rotating component of a gas turbine engine comprises a platform comprising a radially facing endwall and an axial end face, and an airfoil extending radially from the endwall.
  • the end face of the platform faces a seal gap between the rotating component and a stationary component of the gas turbine engine.
  • the seal gap is located between a radially outwardly located hot gas path and a radially inwardly located rotor disc cavity.
  • upstream and downstream positions are defined in relation to a circumferential flow velocity of fluid in the seal gap relative to the end face, resultant from rotation of the rotating component.
  • the end face comprises a groove.
  • the groove comprises a groove floor comprising an inclined portion having increasing depth along a downstream direction, and a sidewall located downstream of the groove floor and facing the groove floor.
  • the sidewall intersects the groove floor extending orthogonal to the end face.
  • a radially outer end of the sidewall is circumferentially offset from a radially inner end of the sidewall, the radially outer end being located downstream of the radially inner end.
  • a seal assembly between a hot gas path and a rotor disc cavity in a gas turbine engine comprises a first seal face formed by an axial end face of a blade platform. From the blade platform, a plurality of blade airfoils extend radially to form a blade assembly. The seal assembly further comprises a second seal face formed by an axial end face of a vane platform. From the vane platform, a plurality of vane airfoils extend radially to form a vane assembly axially adjacent to the blade assembly. The first and second seal faces face each other with a seal gap defined therebetween.
  • the seal gap is located between a radially outwardly located hot gas path and a radially inwardly located rotor disc cavity.
  • upstream and downstream positions are defined in relation to a circumferential flow velocity of fluid in the seal gap relative to the first seal face, resultant from rotation of the blade assembly.
  • the first seal face comprises a plurality of circumferentially spaced grooves.
  • Each groove comprises a groove floor comprising an inclined portion having increasing depth along a downstream direction, and a sidewall located downstream of the groove floor and facing the groove floor.
  • the sidewall intersects the groove floor extending orthogonal to the first seal face.
  • a radially outer end of the sidewall is circumferentially offset from a radially inner end of the sidewall, the radially outer end being located downstream of the radially inner end.
  • FIG. 1 is a diagrammatic sectional view of a portion of a turbine stage of a gas turbine engine including a seal assembly in accordance with an embodiment of the present invention
  • FIG. 2 illustrates an axial end of a rotating blade assembly having pumping features embodied as grooves in accordance with an embodiment of the present invention
  • FIG. 3 illustrates a perspective view of an end face with a groove in accordance with an embodiment of the present invention.
  • FIG. 4 - 7 schematically illustrate, in axial end view, various configurations of the groove in accordance with aspects of the present invention.
  • FIG. 1 a portion of a gas turbine engine 10 is illustrated diagrammatically including a stationary vane assembly 12 and a rotating blade assembly 18.
  • the vane assembly 12 includes a plurality of vane airfoils 14 arranged in a circumferential row. Each vane airfoil 14 is mounted between an inner diameter vane platform or shroud 80 and an outer diameter vane platform or shroud (not shown). As illustrated, the vane platform 80 comprises a radially facing endwall 82 from which one or more vane airfoils 14 extend radially outward.
  • the blade assembly 18 includes a plurality of blades 20 arranged in a circumferential row. Each blade airfoil 20 is mounted on a blade platform 40.
  • the blade platform 40 is located at a radially inner end of the blade airfoil(s) 20 and comprises a radially facing endwall 42 from which one or more blade airfoils 20 extend radially outward.
  • the blade platform 40 is mounted on a rotor disc structure 22 that forms part of a turbine rotor 24.
  • the vane assembly 12 and the blade assembly 18 may be collectively referred to herein as a "stage" of a turbine section of the engine 10, which may include a plurality of stages as will be apparent to those having ordinary skill in the art.
  • the blade assembly may include an outer diameter blade platform, commonly referred to as "tip shroud" at a radially outer end of the blade airfoils 20.
  • the vane assemblies 12 and blade assemblies 18 are spaced apart from one another in an axial direction defined along a longitudinal axis LA of the engine 10, wherein the vane assembly 12 illustrated in FIG. 1 is upstream from the illustrated blade assembly 18 with respect to a flow of hot working gas !3 ⁇ 4.
  • the vane airfoils 14 and the blade airfoils 20 extend into an annular hot gas path 34 defined within the turbine section.
  • the hot working gas HG comprising hot combustion gases is directed through the hot gas path 34 and flows past the vane airfoils 14 and the blade airfoils 20 to remaining stages during operation of the engine 10. Passage of the working gas 3 ⁇ 4 through the hot gas path 34 causes rotation of the blade airfoils 20 and the corresponding blade assembly 18 to provide rotation of the turbine rotor 24.
  • a disc cavity 36 is located radially inwardly from the hot gas path 34 between the ID vane platform 80 and the rotor disc structure 22.
  • Purge air PA such as, for example, compressor bleed air, is provided into the rotor disc cavity 36 to cool the vane platform 80, the blade platform 40 and the rotor disc structure 22.
  • the purge air PA also provides a pressure balance against the pressure of the working gas 3 ⁇ 4 flowing through the hot gas path 34 to counteract an ingestion of the working gas 3 ⁇ 4 into the rotor disc cavity 36.
  • the purge air PA may be provided to the disc cavity 36 from cooling passages (not shown) formed through the rotor 24 and/or from other upstream passages (not shown) as desired. It is noted that additional disc cavities (not shown) are typically provided between remaining ID vane platforms 80 and corresponding adjacent rotor disc structures 22.
  • Embodiments of the present invention provide a seal assembly 100 between a hot gas path 34 and a rotor disc cavity 36.
  • the seal assembly 100 is defined by a first seal face 44, which is formed by an axial end face 44 of the blade platform 40, and a second seal face 84, which is formed by an axial end face 84 of the vane platform 80.
  • the first and second seal faces 44, 84 face each other, with a seal gap 50 being defined therebetween.
  • the seal gap 50 is located between a radially outwardly located hot gas path 34 and a radially inwardly located rotor disc cavity 36.
  • the rotating first seal face 44 is provided with pumping features, embodied as grooves 60.
  • FIG. 2 depicts an axial end of a blade assembly 18 incorporating pumping features in accordance with a first embodiment of the invention.
  • a blade assembly 18 includes a blade platform 40 on which one or more blade airfoils 20 are mounted.
  • Each blade airfoil 20 comprises an aerodynamically shaped outer wall, having a generally concave shaped pressure side 72 and a generally convex shaped suction side 74.
  • the pressure side 72 and the suction side 74 are joined at a leading edge 76 and at a trailing edge 78, which form axial ends of the blade airfoil 20.
  • the blade airfoils 20 extend radially outward from an endwall 42 of the platform 40 into the hot gas path.
  • An axial end of the blade platform 40 is defined by an end face 44.
  • the end face 44 intersects the endwall 42 at a platform edge 45, and extends radially inward from said platform edge 45.
  • the end face 42 may be parallel to the radial direction or be inclined with respect to the radial direction.
  • the illustrated end face 44 is located at a forward axial end 47 of the blade platform 40.
  • a corresponding end face (not shown) may be located at an aft axial end 49 of the blade platform 40.
  • rotor side pumping features are provided on the forward end face 44 of the blade platform 40.
  • the forward end face 44 of the blade platform 40 forms the first seal face of the illustrated seal assembly 100 (see FIG. 1).
  • the second seal face is formed by the aft end face 84 of the ID vane platform 80 of the axially upstream vane assembly 12.
  • pumping features may be formed on the aft end face (not shown) of the blade platform 40.
  • a corresponding seal assembly may be defined by the aft end face of the blade platform 40 and a forward end face of the ID vane platform of an axially downstream vane assembly (not shown).
  • the illustrated pumping features may be provided on a forward or aft end face of an OD blade platform (i.e., tip shroud).
  • a seal assembly may be defined by the forward or aft end face of the OD blade platform and an aft or forward end face of the OD vane platform of an axially adjacent vane assembly.
  • FIG. 3 illustrates a perspective view of the forward end face 44 of the blade assembly comprising rotor side pumping features.
  • the depiction of the blade airfoils is omitted in FIG. 3.
  • the rotor side pumping features comprise a plurality of circumferentially spaced grooves 60 provided on the end face 44 of the blade platform 40.
  • the rotation direction of the blade assembly 18 is indicated by the arrow R, while the circumferential fluid flow velocity in the seal gap relative to the end face 44 is indicated by the arrow F (see FIG. 2).
  • each groove 60 comprises a groove floor 62 and a sidewall 64 located downstream of the groove floor 62.
  • the groove floor 62 defines a recess having a depth in a direction generally perpendicular to the end face 44.
  • the groove floor 62 includes at least an inclined portion 62a having increasing depth along the downstream direction.
  • the groove floor 62 further comprises a flat portion 62b at a constant depth.
  • the flat portion 62b is located between the inclined portion 62a of the groove floor 62 and the sidewall 64.
  • the sidewall 64 faces the groove floor 62 and extends orthogonal to the end face, intersecting the groove floor 62 at an edge 63.
  • the edge 63 is adjacent to the flat portion 62b of the groove floor.
  • the groove floor 62 and the sidewall 64 are thus configured to function as a scooper which faces the rotation of the fluid in the seal gap, pushing said fluid into the hot gas path.
  • the sidewall 64 may be arc-shaped, having a concave side facing the incident seal gap fluid (see FIG. 2 and 4).
  • the pumping action of the rotating groove imparts additional circumferential momentum to the seal gap fluid by accelerating the tangential component of the flow vector of the seal gap fluid, thereby minimizing aerodynamic losses due to the mixing of the seal gap fluid with the fluid in the hot gas path.
  • the sidewall 64 of the groove 60 extends non-parallel to the radial direction, from a radially inner end 66 of the sidewall 64 to a radially outer end 66 of the sidewall 64. That is, the radially outer end 68 of the sidewall 64 is circumferentially offset from the radially inner end 66 of the sidewall 64. In particular, the radially outer end 68 is located circumferentially downstream of the radially inner end 66.
  • the above-illustrated configuration of the sidewall 64 ensures that the momentum imparted to the seal gap fluid by the rotating groove also has a component in the radially outward direction. On account of the increased radial momentum, the purge air functions as an invisible curtain minimizing hot gas ingestion into the rotor disc cavity 36 (see FIG. 1).
  • turbine stage efficiency may be increased.
  • the circumferential locations of the grooves 60 on the end face 44 are near respective leadings edges 76 of the blade airfoils 20. It has been observed that horse-shoe vortices are normally formed on the platform endwall 42 adjacent to the leading edges 76 of the airfoil 20. The high pressure arising out of such horse-shoe vortices tends to discourage the seal gap fluid from flowing out of the rotor disc cavity 36 into the hot has path 34.
  • grooves 60 in circumferential locations near the leading edges 76 would provide the benefit of altering the pressure differential between the horse-shoe vortices and the rotor disc cavity 36, to facilitate a smoother outflow of the seal gap fluid. It should be understood that the above illustrated circumferential placement of the grooves 60 is merely exemplary and other circumferential locations for the grooves 60 may be considered.
  • FIG. 4 - 7 schematically illustrate various configurations of the groove in accordance with aspects of the present invention.
  • the sidewall 64 has a curved profile extending from the radially inner end 66 to the radially outer end 68, having a concave face adjacent to the groove floor 62 which faces the seal gap fluid in the groove.
  • the inclined portion 62a of the groove floor 62 has a decreasing radial width W R in the circumferentially upstream direction. The reduction of the flow cross-section in the circumferential direction serves to accelerate the seal gap fluid which is pushed tangentially by the arc-shaped sidewall 64 in the direction of rotation R.
  • FIG. 4 schematically illustrate various configurations of the groove in accordance with aspects of the present invention.
  • the sidewall 64 has a curved profile extending from the radially inner end 66 to the radially outer end 68, having a concave face adjacent to the groove floor 62 which faces the seal gap fluid in the groove.
  • the inclined portion 62a of the groove floor 62 is triangular shaped.
  • other shapes may be employed for the inclined portion 62a.
  • the inclined portion 62a of the groove floor 62 may have a semi-circular or semi-elliptical shape.
  • the sidewall 64 may be planar, i.e., having a linear profile extending from the radially inner end 66 to the radially outer end 68, as shown in FIG. 6.
  • the intermediate flat portion of the groove may be eliminated, whereby the sidewall 64 intersects directly with the inclined portion 62a, as shown in FIG. 7.

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

Abstract

Un composant rotatif de turbine (18) comprend un profil aérodynamique (20) s'étendant radialement à partir d'une plateforme (40). Une face d'extrémité axiale (44) de la plateforme (40) fait face à un espace d'étanchéité (50) entre le composant rotatif (18) et un composant fixe (12). L'espace d'étanchéité (50) est situé entre un trajet de gaz chaud radialement externe (34) et une cavité de disque de rotor radialement interne (36). Le long de la face d'extrémité (44), des positions amont et aval sont définies par rapport à une vitesse d'écoulement circonférentielle (F) de fluide dans l'espace d'étanchéité (50) par rapport à la face d'extrémité (44), résultant de la rotation du composant rotatif (18). La face d'extrémité (44) comprend une rainure (60). La rainure (60) a un fond de rainure (62) comprenant une partie inclinée (62a) ayant une profondeur croissante le long d'une direction aval, et une paroi latérale (64) en aval du fond de rainure (62) et faisant face au fond de rainure (62). La paroi latérale (64) coupe le fond de rainure (62) s'étendant orthogonalement à la face d'extrémité (44). Une extrémité radialement externe (68) de la paroi latérale (64) est située en aval d'une extrémité radialement interne (66) de celle-ci.
PCT/US2017/012262 2017-01-05 2017-01-05 Ensemble joint d'étanchéité entre un trajet de gaz chaud et une cavité de disque de rotor WO2018128609A1 (fr)

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PCT/US2017/012262 WO2018128609A1 (fr) 2017-01-05 2017-01-05 Ensemble joint d'étanchéité entre un trajet de gaz chaud et une cavité de disque de rotor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2017/012262 WO2018128609A1 (fr) 2017-01-05 2017-01-05 Ensemble joint d'étanchéité entre un trajet de gaz chaud et une cavité de disque de rotor

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11286784B2 (en) 2020-02-13 2022-03-29 Rolls-Royce Plc Aerofoil assembly and method
US11371356B2 (en) 2020-02-13 2022-06-28 Rolls-Royce Plc Aerofoil assembly and method

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2004036510A (ja) * 2002-07-04 2004-02-05 Mitsubishi Heavy Ind Ltd ガスタービン動翼シュラウド
US20130017080A1 (en) * 2011-07-12 2013-01-17 Kok-Mun Tham Flow directing member for gas turbine engine
US20140205443A1 (en) * 2013-01-23 2014-07-24 Siemens Aktiengesellschaft Seal assembly including grooves in an aft facing side of a platform in a gas turbine engine
EP3064709A1 (fr) * 2015-03-02 2016-09-07 General Electric Company Plate-forme d'aube de turbine pour influencer les pertes d'incursion de gaz chaude

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2004036510A (ja) * 2002-07-04 2004-02-05 Mitsubishi Heavy Ind Ltd ガスタービン動翼シュラウド
US20130017080A1 (en) * 2011-07-12 2013-01-17 Kok-Mun Tham Flow directing member for gas turbine engine
US20140205443A1 (en) * 2013-01-23 2014-07-24 Siemens Aktiengesellschaft Seal assembly including grooves in an aft facing side of a platform in a gas turbine engine
EP3064709A1 (fr) * 2015-03-02 2016-09-07 General Electric Company Plate-forme d'aube de turbine pour influencer les pertes d'incursion de gaz chaude

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11286784B2 (en) 2020-02-13 2022-03-29 Rolls-Royce Plc Aerofoil assembly and method
US11371356B2 (en) 2020-02-13 2022-06-28 Rolls-Royce Plc Aerofoil assembly and method

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