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1. (WO2018128609) SEAL ASSEMBLY BETWEEN A HOT GAS PATH AND A ROTOR DISC CAVITY
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SEAL ASSEMBLY BETWEEN A HOT GAS PATH AND A ROTOR DISC

CAVITY

BACKGROUND

1. Field

[0001] 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.

2. Description of the Related Art

[0002] 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. Within the turbine section are 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.

[0003] 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. Such 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.

[0004] To counteract the hot gas ingestion and to aid in cooling the blade root core, 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.

SUMMARY

[0005] Briefly, 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.

[0006] According a first aspect of the present invention, a rotating component of a gas turbine engine is provided. The rotating component 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. Along the end face, 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.

[0007] According to a second aspect of the present invention, a seal assembly between a hot gas path and a rotor disc cavity in a gas turbine engine is provided. The seal assembly 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. Along the first seal face, 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.

BRIEF DESCRIPTION OF THE DRAWINGS

[0008] The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.

[0009] 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;

[0010] 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;

[0011] FIG. 3 illustrates a perspective view of an end face with a groove in accordance with an embodiment of the present invention; and

[0012] FIG. 4 - 7 schematically illustrate, in axial end view, various configurations of the groove in accordance with aspects of the present invention.

DETAILED DESCRIPTION

[0013] In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

[0014] Referring to 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. As illustrated, 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. In some turbine stages, particularly in low pressure turbine stages, 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 !¾.

[0015] 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 ¾ 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.

[0016] Referring to FIG. 1, 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 ¾ flowing through the hot gas path 34 to counteract an ingestion of the working gas ¾ 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.

[0017] 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. In accordance with aspects of the present invention, the rotating first seal face 44 is provided with pumping features, embodied as grooves 60. The grooves 60 are geometrically configured to impart a circumferential as well as a radially outward momentum to the fluid flow in the seal gap 50. The radially outward component of the imparted momentum assists in minimizing hot gas ingestion into the rotor disc cavity 36, while the circumferential component of the imparted momentum assists in minimizing aerodynamic losses when the purge air PA mixes with the hot working gas HG. Exemplary configurations of the grooves are illustrated below referring to FIG. 2 [0018] FIG. 2 depicts an axial end of a blade assembly 18 incorporating pumping features in accordance with a first embodiment of the invention. As shown, 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.

[0019] In the example shown in FIG. 2, 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). In the illustrated example, 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. Although not illustrated in the drawings, it is to be understood that additionally or alternately, such pumping features may be formed on the aft end face (not shown) of the blade platform 40. In such a case, 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). As a further variant, in case of a shrouded blade, the illustrated pumping features may be provided on a forward or aft end face of an OD blade platform (i.e., tip shroud). In this case, 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.

[0020] FIG. 3 illustrates a perspective view of the forward end face 44 of the blade assembly comprising rotor side pumping features. For the sake of simplicity, the depiction of the blade airfoils is omitted in FIG. 3. Referring jointly to FIG. 2 and 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). Along the end face 44, upstream and downstream positions may be thereby defined in relation to the circumferential flow velocity F of fluid in the seal gap relative to the end face 44, resultant from rotation of the blade assembly 18. The seal gap fluid comprises purge air, and may also comprise some amount of the hot working gas. As shown in FIG. 2 and 3, 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. In particular, the groove floor 62 includes at least an inclined portion 62a having increasing depth along the downstream direction. In the illustrated embodiment, 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. In this example, 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. In order to enhance the scooping effect, 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.

[0021] In accordance with the illustrated embodiments, 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).

[0022] On account of the decreased aerodynamic mixing losses and reduced hot gas ingestion, turbine stage efficiency may be increased. In the illustrated non-limiting example (see FIG. 2), 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. Placing the 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.

[0023] FIG. 4 - 7 schematically illustrate various configurations of the groove in accordance with aspects of the present invention. In the embodiment of FIG. 4, 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. Furthermore, as shown in FIG. 4, the inclined portion 62a of the groove floor 62 has a decreasing radial width WR 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. In the example of FIG. 4, the inclined portion 62a of the groove floor 62 is triangular shaped. However, other shapes may be employed for the inclined portion 62a. For example, as shown in FIG. 5, the inclined portion 62a of the groove floor 62 may have a semi-circular or semi-elliptical shape. In another variant, instead of having a curved profile, 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. In yet another variant, 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.

[0024] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.