US20180223683A1

US20180223683A1 - Gas turbine seal arrangement

Info

Publication number: US20180223683A1
Application number: US15/743,117
Authority: US
Inventors: Kok-Mun Tham; Abdullatif M. Chehab; Patrick M. Pilapil; Yan Yin; Christian Xavier Campbell; Vincent Paul Laurello
Original assignee: Siemens Energy Inc
Current assignee: Siemens Energy Inc
Priority date: 2015-07-20
Filing date: 2015-07-20
Publication date: 2018-08-09
Also published as: WO2017014737A1; EP3325779A1

Abstract

A turbine arrangement including a rotor and a stator surrounding the rotor and comprising guide vane segments, each guide vane segment comprising an airfoil and a radially inner vane platform. A seal arrangement includes a static seal inward from the inner vane platforms and having a radially extending face plate, first and second cylindrical seal walls extending from outer and inner ends of the annular face plate, an annular seal plate extending radially from the second cylindrical seal wall, and an angel wing extending between the first cylindrical seal wall and the annular seal plate to define a first annular cavity and a second annular cavity. Circumferentially spaced cut-outs define passages through the annular seal plate between the first and second annular cavities and are aligned with fasteners that attach the annular face plate to a support ring for supporting the inner vane platform.

Description

FIELD OF THE INVENTION

The present invention relates to gas turbine engines and, in particular, to seal arrangements providing a seal between a hot gas flow path and a disk cavity supplied with secondary air.

BACKGROUND OF THE INVENTION

In a gas turbine engine, hot combustion gases are routed from a combustor to a turbine section, in which stator vanes are designed to direct the hot gases onto rotor blades resulting in rotational movement of a rotor to which the rotor blades are connected. Radially inwards and outwards of airfoils of these stator vanes and rotor blades, platforms, casing structure, or other components may be present such as to form an annular fluid passage into which the airfoils of the stator vanes and the rotor blades extend and through which the hot combustion gases pass.
As the rotating rows of rotor blades and non-rotating rows of stator vanes are arranged alternately, gaps may be present between the rows of rotor blades and the rows of stator vanes. Seal structure is typically provided to reduce the size of the gaps and/or to seal these gaps so as to minimize or limit the amount of hot combustion gas that is lost via these gaps and to minimize the amount of secondary air that can pass into the hot gas flow. The structure to seal these gaps between rotor blades and stator vanes is commonly referred to as a turbine rim seal.
In the turbine front stages, effective operation of the rim seal is particularly important to ensure mechanical integrity of the steel turbine disks, as even a small amount of ingestion of gas from the gas path can potentially raise turbine disk cavity temperatures significantly. In turbine engines where the first row turbine blade platforms are cooled by air supplied from the first disk cavity, an effective rim seal ensures effectual blade platform cooling, and has a reduced requirement for cavity purge flow. Hence, an improvement in rim sealing can result in a reduction in overall cooling and sealing air consumption, and an improved turbine aerodynamic performance, since mixing loss from purge flow induction can be reduced. The turbine rim seal on the upstream side of the first row of turbine blades can comprise a stationary static seal housing a honeycomb for mating with a rotor angel wing. The static seal may be held in position by bolts having heads that extend into the disk cavity and which can increase drag in the disk cavity, leading to increased cavity temperatures due to windage.

SUMMARY OF THE INVENTION

In accordance with an aspect of the invention, a turbine arrangement is provided comprising a rotor that rotates about a rotor axis and comprises a plurality of rotor blade segments extending radially outward, each rotor blade segment comprises an airfoil and a radially inner blade platform. A stator surrounds the rotor so as to form an annular flow path for a hot working gas, and the stator comprises a plurality of guide vane segments disposed adjacent the plurality of rotor blade segments. The plurality of guide vane segments extend radially inward, each guide vane segment comprising an airfoil and a radially inner vane platform. A seal arrangement comprises an annular face plate extending radially inward from the vane platform, a first cylindrical seal wall extending axially from an outer end of the face plate, a second cylindrical seal wall extending axially from an inner end of the face plate, an annular seal plate extending radially from an end of the second cylindrical seal wall, and an angel wing extending from the rotor and having a distal end between the first cylindrical seal wall and the seal plate to define a first annular cavity and a second annular cavity. The first annular cavity is defined at least by the first and second cylindrical seal walls and the annular seal plate. The second annular cavity is defined at least by the angel wing and the annular seal plate. The first annular cavity is in fluid communication with the annular flow path via a first annular seal passage between the first cylindrical seal wall and the angel wing. The first annular cavity is in fluid communication with the second annular cavity via a second annular seal passage between the angel wing and an outer end of the annular seal plate. The annular face plate is attached to a support ring that supports the inner vane platform, including a plurality of circumferentially spaced fasteners passing through apertures in the annular face plate into the support ring. A plurality of circumferentially spaced cut-outs are formed in the annular seal plate defining passages between the first and second annular cavities.
The fasteners may include fastener heads that are located in the first annular cavity. The cut-outs in the annular seal plate may each be circumferentially aligned with a fastener head. The cut-outs can each be defined by a sidewall that is angled circumferentially in a direction of rotor rotation extending from the second annular cavity toward the first annular cavity.
A cylindrical flange may extend parallel to the first and second cylindrical seal walls into the first annular cavity from the outer end of the annular seal plate.
A surface at the outer end of the annular seal plate can be angled radially inward from the first cavity toward the second cavity.
The surface at the outer end of the annular seal plate can be defined by an inner seal member affixed to the cylindrical flange and cooperate with the angel wing to define the second annular seal passage.
The surface at the outer end of the annular seal plate can be stepped radially inward from the first cavity toward the second cavity.
An outer inner seal member can be affixed to an inner side of the first cylindrical seal wall and cooperate with the angel wing to define the first annular seal passage.
The distal end of the angel wing can be formed with a hammerhead configuration cooperating with surfaces at the first cylindrical seal wall and the outer end of the annular seal plate to define the first and second annular seal passages, respectively.
The outer end of the annular seal plate can be formed with a knife-edge and can cooperate with the angel wing to define the second annular seal passage.
A knife-edge can extend radially inward from the first cylindrical seal wall and can cooperate with the angel wing to define the first annular seal passage.
In accordance with another aspect of the invention, a turbine arrangement is provided comprising a rotor that rotates about a rotor axis and comprises a plurality of rotor blade segments extending radially outward, each rotor blade segment comprises an airfoil and a radially inner blade platform. A stator surrounds the rotor so as to form an annular flow path for a hot working gas, and the stator comprises a plurality of guide vane segments disposed adjacent the plurality of rotor blade segments. The plurality of guide vane segments extend radially inward, each guide vane segment comprising an airfoil and a radially inner vane platform. A seal arrangement comprises an annular face plate extending radially inward from the vane platform, a first cylindrical seal wall extending axially from an outer end of the face plate, a second cylindrical seal wall extending axially from an inner end of the face plate, an annular seal plate extending radially from an end of the second cylindrical seal wall, and an angel wing extending from the rotor and having a distal end between the first cylindrical seal wall and the seal plate to define a first annular cavity and a second annular cavity. The first annular cavity is defined at least by the first and second cylindrical seal walls and the annular seal plate. The second annular cavity is defined at least by the angel wing and the annular seal plate. The first annular cavity is in fluid communication with the annular flow path via a first annular seal passage between the first cylindrical seal wall and the angel wing. The first annular cavity is in fluid communication with the second annular cavity via a second annular seal passage between the angel wing and an outer end of the annular seal plate. The annular face plate is attached to a support ring that supports the inner vane platform, including a plurality of circumferentially spaced fasteners passing through apertures in the annular face plate into the support ring. A plurality of circumferentially spaced cut-outs in the annular seal plate define passages between the first and second annular cavities. The cut-outs are each circumferentially aligned with a fastener and are defined by a sidewall that is angled circumferentially in a direction of rotor rotation extending from the second annular cavity toward the first annular cavity.
A cylindrical flange may extend parallel to the first and second cylindrical seal walls into the first annular cavity from the outer end of the annular seal plate. An inner seal member can be affixed to the cylindrical flange and cooperate with the angel wing to define the second annular seal passage. The inner seal member can have an outer sealing surface that has a reduced downstream radial dimension adjacent to the second annular cavity in comparison to the upstream radial dimension of the inner seal member adjacent to the first annular cavity.
In accordance with a further aspect of the invention, a turbine arrangement is provided comprising a rotor that rotates about a rotor axis and comprises a plurality of rotor blade segments extending radially outward, each rotor blade segment comprises an airfoil and a radially inner blade platform. A stator surrounds the rotor so as to form an annular flow path for a hot working gas, and the stator comprises a plurality of guide vane segments disposed adjacent the plurality of rotor blade segments. The plurality of guide vane segments extend radially inward, each guide vane segment comprising an airfoil and a radially inner vane platform. A seal arrangement comprises an annular face plate extending radially inward from the vane platform, a first cylindrical seal wall extending axially from an outer end of the face plate, a second cylindrical seal wall extending axially from an inner end of the face plate, an annular seal plate extending radially from an end of the second cylindrical seal wall, and an angel wing extending from the rotor and having a distal end between the first cylindrical seal wall and the seal plate to define a first annular cavity and a second annular cavity. The first annular cavity is defined at least by the first and second cylindrical seal walls and the annular seal plate. The second annular cavity is defined at least by the angel wing and the annular seal plate. The first annular cavity is in fluid communication with the annular flow path via a first annular seal passage between the first cylindrical seal wall and the angel wing. The first annular cavity is in fluid communication with the second annular cavity via a second annular seal passage between the angel wing and an outer end of the annular seal plate. A cylindrical flange extends parallel to the first and second cylindrical seal walls into the first annular cavity from the outer end of the annular seal plate.
A plurality of circumferentially spaced cut-outs in the annular seal plate may define passages between the first and second annular cavities. The annular face plate may be attached to a support ring that supports the inner vane platform, and a plurality of circumferentially spaced fasteners can pass through apertures in the annular face plate into the support ring and may be located in circumferential alignment with the cut-outs.
An inner seal member can be affixed to the cylindrical flange and can cooperate with the angel wing to define the second annular seal passage.

BRIEF DESCRIPTION OF THE DRAWINGS

While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:

FIG. 1 is a schematic section through a turbine section of a gas turbine engine illustrating a seal arrangement in accordance with an aspect of the present invention;

FIG. 2 is an axial view of a portion of an annular seal plate in a seal arrangement illustrating an aspect of the present invention;

FIG. 2A is a cross-sectional view taken along line 2A-2A in FIG. 2;

FIG. 3A is a schematic section similar to FIG. 1 illustrating a variation of the seal arrangement;

FIG. 3B is a schematic section similar to FIG. 1 illustrating a variation of the seal arrangement;

FIG. 4 is a schematic section similar to FIG. 1 illustrating a variation of the seal arrangement;

FIG. 5 is a schematic section similar to FIG. 1 illustrating a variation of the seal arrangement;

FIG. 6 is a schematic section similar to FIG. 1 illustrating a variation of the seal arrangement; and

FIG. 7 is a schematic section through a turbine section of a prior art gas turbine engine.

DETAILED DESCRIPTION OF THE INVENTION

In the following detailed description of the preferred embodiment, 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 preferred 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.
The present invention is directed to a turbine arrangement such as may comprise a gas turbine engine comprising a compressor section, a combustor section and a turbine section which are arranged adjacent to each other. In operation of the gas turbine engine, ambient air may be compressed by the compressor section, mainly provided as an input to the combustor section with one or more combustors. In the combustor section the compressed air can be mixed with liquid and/or gaseous fuel and this mixed fluid is burnt, resulting in a hot working gas. The hot working gas is then guided from the combustor to the turbine section, in which the hot working gas will drive one or more rows of rotor blades resulting in a rotational movement of a shaft.
The direction of the fluid flow will be called “downstream” from the inlet via the compressor section, via the combustor section to the turbine section and finally to an exhaust. The opposite direction will be called “upstream”. The term “leading” corresponds to an upstream location, “trailing” corresponds to a downstream location. The turbine section may be substantially rotational symmetric about an axis of rotation. A positive axial direction may be defined as the downstream direction. In the figures provided herein, the hot working gas will be guided substantially from left to right in parallel to the positive axial direction.
Referring now to FIG. 1, a set of guide vanes 10 and rotor blades 12 are shown, it being understood that the guide vanes 10 and rotor blades 12 are located in respective circumferentially extending rows about a rotational center axis A_c. The first set of guide vanes 10 is located immediately downstream of the combustor section (not shown). Each guide vane 10 in the set of guide vanes 10 includes an airfoil 14 extending in an approximately radial direction, indicated by arrow r, with respect to the center axis A_cof the turbine section and an outer platform (not shown) for the mounting of the guide vane 10 in a housing or a casing, the housing and the outer platform being a part of a stator, i.e. being non-rotational. Each airfoil 14 extends radially inward from the outer platform to an inner vane platform 16 of the guide vane 10 for forming a stationary, annular supporting structure at a radially inner position of the airfoils 14 of the guide vane 10.
The outer platform, inner vane platform 16 and the airfoil 14 typically are built as a one-piece guide vane segment and a plurality of guide vane segments are arranged circumferentially around the center axis A_cto build one guide vane stage, and is generally referred to as the stator 17. The outer platform and inner platform 16 are arranged to form an annular flow path or flow passage 18 for hot working gases to flow in the flow direction, indicated by an arrow with reference sign 20. Consequently, the outer platforms and inner platforms 16 may need to be cooled, such as by cooling air provided directly from the compressor section of the gas turbine engine without passing through combustors in the combustion section.
Immediately downstream of the illustrated guide vane stage, there is the first rotor stage including a number of rotor blades 12. The rotor blades 12 comprise an inner platform 22 and an outer shroud (not shown) forming a continuation of the annular flow path 18 so that the hot working gas will be guided downstream as indicated by arrow a (or arrow with reference symbol 20). A plurality of rotor blades 12 extend outward between the inner platform 22 and the outer shroud. A single inner platform section 22 and a single rotor blade airfoil 24 may form one rotor blade segment. A plurality of rotor blade segments are connected to a rotor disc 26 supported for rotational movement and defining a portion of a rotor shaft, the assembled structure being generally referred to as a rotor 28.
In accordance with an aspect of the invention, a seal arrangement 30 is provided between the rotating parts, i.e., the rotor 28, and the stationary parts, i.e., the stator 17, so that the hot working gas will stay in the annular flow path 18 and will not mix directly with a secondary fluid, e.g., air provided for cooling. The seal arrangement 30 will be described herein with reference to a location between the row 1 vanes and the row 1 blades forming a first turbine stage, however, it may be understood that the concept described herein may be incorporation at other locations including between adjacent vanes and blades of other stages in the turbine section.
Referring now to FIG. 7, a prior art turbine arrangement is shown comprising a stator 117 for which a guide vane 110 is shown. The guide vane 110 comprises an outer platform 115, an inner platform 116, and an airfoil 114. Furthermore the turbine arrangement also comprises a rotor 128 for which a rotor blade 112 is shown. The rotor blade 112 comprises an inner blade platform 122 and an airfoil 124. Further, a shroud 113 may be provided at a radial distant end of the rotor blade 112, the distant end being at an opposite end compared to the inner blade platform 122. Between the mentioned outer and inner platforms an annular flow path 118 is formed through which a hot working gas, indicated by an arrow 120, is guided to drive the plurality of rotor blades 112.
A seal arrangement 130 formed according to the prior art is shown between the guide vane 110 and the rotor blade 112. The seal arrangement 130 provides a sealing mechanism between the guide vane 110 and rotor blade 112. Hot gases from the main annular flow path 118 may enter the seal arrangement 130 during operation. In other modes of operation, secondary air 132B may enter the main annular flow path 118. This may be caused by a pressure difference between a provided secondary air 132A and the pressurized hot working gas 120 in the main annular flow path 118. The pressure difference may be caused by local pressure gradients surrounding the blades and vanes at the seal arrangement 130 during operation of the gas turbine engine.
Referring again to FIG. 1, details of the seal arrangement 30 according to an aspect of the invention will be described. The seal arrangement 30 is depicted located radially inward from a downstream portion of the inner vane platform 16 and upstream from the inner blade platform 22. The seal arrangement 30 comprises a static seal member 31 including an annular face plate 34 extending in a radial plane radially inward from the inner vane platform 16. A first cylindrical seal wall 36 extends axially from an outer end 38 of the face plate 34, and a second cylindrical seal wall 40 extends axially from an inner end 42 of the face plate 34. An annular seal plate 44 extends radially from an axial downstream end 46 of the second cylindrical seal wall 40. The seal arrangement additionally includes an angel wing 48 extending from the rotor 28, i.e., from an axial forward side of the rotor disk 26, and having a distal end 50 between the first cylindrical seal wall 36 and an outer end 52 of the seal plate 44 to define a first annular cavity C₁and a second annular cavity C₂.
The first annular cavity C₁is defined at least by the first and second cylindrical seal walls 36, 40 and the annular seal plate 44, and is further defined by the annular face plate 34. The second annular cavity C₂is defined at least by the angel wing 48 and the annular seal plate 44, and can be further defined by the rotor disk 26, wherein it may be understood that at least a portion of the second annular cavity C₂is radially aligned with the first annular cavity C₁, and is located on an axially opposite side of the annular seal plate 44 from the first annular cavity C₁. Additionally, it may be noted that the second annular cavity C₂corresponds to a disk cavity that receives a supply of secondary air, i.e., cooling and purge air, from the compressor for supplying platform coolant to the platform 22 for the rotor blade 12.
The first annular cavity C₁is in limited fluid communication with the annular flow path 18 via a first annular seal passage P₁between the first cylindrical seal wall 36 and the angel wing 48. In particular, the first annular seal passage P₁can be formed between a radially extending rim portion 50 a defined on the distal end 50 of the angel wing 48 and an outer circumferential seal member 54, such as a honeycomb seal, located on a radial inner side of the inner vane platform 16.
The first annular cavity C₁is in limited fluid communication with the second annular cavity C₂via a second annular seal passage P₂between the angel wing 48 and the outer end 52 of the annular seal plate 44. In particular, the second annular seal passage P₂can be formed between an inner side of the distal end 50 of the angel wing 48 and an inner circumferential seal member 56, such as a honeycomb seal, located on the radial outer end 52 of the annular seal plate 44. In this regard, a cylindrical flange 58 can be formed extending parallel to the first and second cylindrical seal walls 36, 40 from the outer end 52 of the annular seal plate 44 into the first cavity C₁and defines a support surface for the seal member 56.
An axial forward side 60 of the axial distal end 50 a of the angel wing 48 faces toward and cooperates with a surface on the annular face plate 34, which may optionally be provided by a honeycomb seal member 61. In particular, as the gas turbine engine ramps up to a steady state temperature and operating speed, the stator and rotor can shift or move axially and radially relative to each other, such as by movement of the honeycomb seal member 61 toward the axial forward side 60 on the angel wing 48.
The annular face plate 34 is attached to a support ring 62 that supports the inner vane platform 16. The support ring 62 can be conventional stationary vane support structure on the interior of the turbine assembly and may be supported, for example, to a compressor discharge casing (not shown). The annular face plate 34 may include a planar face surface 34 a that is in facing engagement with a planar facing surface 62 a of the support ring 62. The annular face plate 34 can be rigidly affixed to the support ring 62 by a plurality of circumferentially spaced fasteners 64, such as bolts, passing through apertures 66 in the annular face plate 34 into the support ring 62. The fasteners 64 can typically include fastener or bolt heads 64 a that extend from the annular face plate 34 into the first annular cavity C₁.
Referring to FIG. 2, the annular seal plate 44 can be provided with a plurality of circumferentially spaced cut-outs 68 defining passages between the first and second annular cavities C₁, C₂. In particular, the cut-outs 68 can extend radially inward through the outer end 52 of the annular seal plate 44 and through a portion of the cylindrical flange 58. The cut-outs 68 are circumferentially aligned with the bolt heads 64 a and provide an access opening for a tool to pass axially through the annular seal plate 44 into engagement with the bolt heads 64 a for mounting and removal of the static seal member 31 to and from the support ring 62. Referring further to FIG. 2A, the cut-outs 68 can be defined by opposing cut-out side walls 68 a, 68 b that are angled circumferentially with respect to the axial direction, as depicted by arrow a in FIG. 1, and as is discussed further below. Additionally, slits 70 may be formed in the annular seal plate 44, extending radially inward from the cut-outs 68 to a location adjacent to the second cylindrical seal wall 40. The slits 70 can optionally be included to provide stress relief to the annular seal plate 44.
Operation of the seal assembly 30 will now be described with respect to operation of the gas turbine engine. As described above, the distal end of the angel wing 48 is positioned in the space between the first cylindrical seal wall 36 and the annular seal plate 44 and rotates relative to the static seal member 31 as the rotor 28 rotates during operation of the engine. The first annular cavity C₁serves as a buffer cavity separating the hot gas flow 20 from the secondary air contained in the disk cavity defined by the second annular cavity C₂. In addition to trapping any hot gas that passes through the first annular seal passage P₁, the first annular cavity C₁damps out any remaining pressure asymmetry associated with pressure in the hot gas path 18 driving ingestion of the hot gases toward the second annular cavity C₂. The cylindrical flange 58, in addition to providing a support surface for the inner seal member 56, also operates to orient flow away from the second annular seal passage P₂, as shown by arrow F₁(FIG. 1), to inhibit hot gas flow entering the first annular cavity C₁from passing into the second annular cavity C₂. That is, the cylindrical flange 58 can direct the isolated flow in the first annular cavity C₁in the upstream direction away from the second annular cavity C₂to limit passage of the hot gases to the second annular cavity C₂. In addition, the junction 59 between the annular seal plate 44 and the cylindrical flange 58 may be rounded to make use of disk pumping flow, i.e., highly-swirled flow from the second annular cavity C₂induced by rotor 28 rotation flowing in axially-upstream direction through P₂to annular cavity C₁, to counter any ingestion flow through the second annular passage P₂from the first annular cavity C₁to the second annular cavity C₂.
The presence of the bolt heads 64 a extending into the first annular cavity C₁further operates to decrease passage of the hot gases into the second annular cavity C₂in that the bolt heads 64 a can increase energy loss of the ingested flow of hot gases, which is highly swirled, reducing the flow energy of the gases trapped in the first annular cavity C₁. The annular seal plate 44 serves as a windage cover to reduce windage in the second annular cavity C₂, such as might otherwise be caused by the bolt heads 64 a as an effect of stationary bolt drag to rotor rotation. Windage can result in heating of the cooling air in the second annular cavity C₂such that the windage cover provided by the annular seal plate 44 can inhibit heating and improve cooling efficiency.
An additional sealing aspect of the seal assembly 30 is provided by the angled sidewalls 68 a, 68 b, as illustrated in FIG. 2A, in that the sidewalls 68 a, 68 b are angled in the downstream circumferential direction of rotor rotation R₁, extending from the second annular cavity C₂toward the first annular cavity C₁. The angled orientation of the sidewalls 68 a, 68 b can operate to use the disk pumping flow, i.e., the circumferential flow in the R₁direction in the second annular cavity C₂, to induce a flow of a portion of secondary air from the second annular cavity C₂toward the first annular cavity C₁to provide some aerodynamic sealing, such as to inhibit flow of the trapped gases from the first annular cavity C₁through the cut-outs 68.
Referring to FIGS. 3A and 3B, a variation on the inner circumferential seal member 56 of FIG. 1 is shown in which the outer sealing surface has a reduced downstream radial dimension in comparison to its upstream radial dimension. FIG. 3A illustrates an inner circumferential seal member 56′ in which its outer sealing surface 56 a′ is ramped or angled radially inward in the downstream direction from the first cavity C₁toward the second cavity C₂. FIG. 3B illustrates an inner circumferential seal member 56″ including an outer dimension that is stepped radially inward from the first annular cavity C₁toward the second annular cavity C₂, and defined by a first outer surface 56 a″ having a greater circumference than a second outer surface 56 b″ that is located stepwise inward from the first outer surface 56 a″. Each of the seal members 56′, 56″ may provide an angled contour that accommodates axial and radial movement of the distal end 50 of the angel wing 48 to maintain a smaller gap at the second annular passage P₂as the gas turbine engine ramps up to steady state temperature and speed. In particular, the configurations shown in FIGS. 3A and 3B are believed to achieve tighter steady state, or hot-running, clearances since it is contoured to align with relative movements of static seal 31 and angel wing 48 expected during transient conditions associated with engine startup.
Referring to FIG. 4, a variation on the outer end 52 of the annular seal plate 44 illustrated in FIG. 1 is shown. FIG. 4 illustrates an outer end 52′ of an annular seal plate 44′ formed with a knife-edge for cooperating with the distal end 50 of the angel wing 48 to define the second annular passage P₂. The knife-edge outer end 52′ is formed with an angled surface 52 a′, angled in the upstream direction and radially outward, that can facilitate disk pumping flow and deter ingestion flow from the first annular cavity C₁to the second annular cavity C₂. Additionally, the knife-edge geometry can minimize damage in the event of unintended contact with the angel wing 48.
FIG. 5 illustrates a further variation similar to FIG. 4 in which the outer circumferential seal member 54 is replaced by an outer annular seal knife-edge 55 extending radially inward from the first cylindrical seal wall 36. The outer knife-edge 55 is formed with an angled surface 55 a, angled in the downstream direction and radially inward to reduce ingestion flow into the first annular cavity C₁, and cooperates with a distal knife-edge end 50′ of the angel wing 48′ to define the first annular passage P₁.
Referring to FIG. 6, a variation on the distal end 50 of the angel wing 48 of FIG. 1 is shown in which a distal end 50″ is formed as a hammerhead having an outwardly extending rim portion 50 a cooperating with the outer seal member 54 and an inwardly extending rim portion 50 b cooperating with the inner seal member 56. The hammerhead distal end 50″ is configured to accommodate the transient clearance behavior resulting in variations in the gaps defined at the first and second annular passages P₁, P₂to provide effective sealing through the entire engine operating cycle. Specifically, during startup, the first annular passage P₁will be tight and rubbing at the seal member 54 may occur, and during steady state operation, the second annular passage P₂will be tight due to the stator 17 being hotter than the rotor 28 and growing or expanding outwardly relative to the rotor 28.
It should be noted that, although the seal arrangement described herein includes reference to honeycomb seals, i.e., seal members 54, 56, 56′, 56″ and 61, these seal members are optional, and the seal arrangement can operate without these seal members or can be provided with other seal elements than the described honeycomb seals.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.

Claims

What is claimed is:

1. A turbine arrangement comprising:

a rotor that rotates about a rotor axis and comprises a plurality of rotor blade segments extending radially outward, each rotor blade segment comprises an airfoil and a radially inner blade platform;

a stator surrounding the rotor so as to form an annular flow path for a hot working gas, the stator comprises a plurality of guide vane segments disposed adjacent the plurality of rotor blade segments, the plurality of guide vane segments extending radially inward, each guide vane segment comprising an airfoil and a radially inner vane platform;

a seal arrangement comprising an annular face plate extending radially inward from the vane platform, a first cylindrical seal wall extending axially from an outer end of the face plate, a second cylindrical seal wall extending axially from an inner end of the face plate, an annular seal plate extending radially from an end of the second cylindrical seal wall, and an angel wing extending from the rotor and having a distal end between the first cylindrical seal wall and the seal plate to define a first annular cavity and a second annular cavity,

wherein:

the first annular cavity is defined at least by the first and second cylindrical seal walls and the annular seal plate;

the second annular cavity is defined at least by the angel wing and the annular seal plate;

the first annular cavity is in fluid communication with the annular flow path via a first annular seal passage between the first cylindrical seal wall and the angel wing;

the first annular cavity is in fluid communication with the second annular cavity via a second annular seal passage between the angel wing and an outer end of the annular seal plate;

the annular face plate is attached to a support ring that supports the inner vane platform, including a plurality of circumferentially spaced fasteners passing through apertures in the annular face plate into the support ring; and

a plurality of circumferentially spaced cut-outs in the annular seal plate defining passages between the first and second annular cavities.

2. The turbine arrangement according to claim 1, characterized in that the fasteners include fastener heads that are located in the first annular cavity.

3. The turbine arrangement according to claim 2, characterized in that the cut-outs in the annular seal plate are each circumferentially aligned with a fastener head.

4. The turbine arrangement according to claim 3, characterized in that the cut-outs are each defined by a sidewall that is angled circumferentially in a direction of rotor rotation extending from the second annular cavity toward the first annular cavity.

5. The turbine arrangement according to claim 1, characterized in that a cylindrical flange extends parallel to the first and second cylindrical seal walls into the first annular cavity from the outer end of the annular seal plate.

6. The turbine arrangement according to claim 1, characterized in that a surface at the outer end of the annular seal plate is angled radially inward from the first cavity toward the second cavity.

7. The turbine arrangement according to claim 6, characterized in that the surface at the outer end of the annular seal plate is defined by an inner seal member affixed to the cylindrical flange and cooperates with the angel wing to define the second annular seal passage.

8. The turbine arrangement according to claim 6, characterized in that the surface at the outer end of the annular seal plate is stepped radially inward from the first cavity toward the second cavity.

9. The turbine arrangement according to claim 6, characterized in that an outer seal member is affixed to an inner side of the first cylindrical seal wall and cooperates with the angel wing to define the first annular seal passage.

10. The turbine arrangement according to claim 1, characterized in that the distal end of the angel wing is formed with a hammerhead configuration cooperating with surfaces at the first cylindrical seal wall and the outer end of the annular seal plate to define the first and second annular seal passages, respectively.

11. The turbine arrangement according to claim 1, characterized in that the outer end of the annular seal plate is formed with a knife-edge and cooperates with the angel wing to define the second annular seal passage.

12. The turbine arrangement according to claim 11, characterized in that a knife-edge extends radially inward from the first cylindrical seal wall and cooperates with the angel wing to define the first annular seal passage.

13. A turbine arrangement comprising:

wherein:

a plurality of circumferentially spaced cut-outs in the annular seal plate defining passages between the first and second annular cavities, wherein the cut-outs are each circumferentially aligned with a fastener and are defined by a sidewall that is angled circumferentially in a direction of rotor rotation extending from the second annular cavity toward the first annular cavity.

14. The turbine arrangement according to claim 13, characterized in that a cylindrical flange extends parallel to the first and second cylindrical seal walls into the first annular cavity from the outer end of the annular seal plate.

15. The turbine arrangement according to claim 14, characterized in that an inner seal member is affixed to the cylindrical flange and cooperates with the angel wing to define the second annular seal passage.

16. The turbine arrangement according to claim 15, characterized in that the inner seal member has an outer sealing surface that has a reduced downstream radial dimension adjacent to the second annular cavity in comparison to the upstream radial dimension of the honeycomb seal adjacent to the first annular cavity.

17. A turbine arrangement comprising:

wherein:

the first annular cavity is in fluid communication with the second annular cavity via a second annular seal passage between the angel wing and an outer end of the annular seal plate; and

a cylindrical flange extends parallel to the first and second cylindrical seal walls into the first annular cavity from the outer end of the annular seal plate.

18. The turbine arrangement according to claim 17, characterized in that a plurality of circumferentially spaced cut-outs in the annular seal plate define passages between the first and second annular cavities.

19. The turbine arrangement according to claim 18, characterized in that the annular face plate is attached to a support ring that supports the inner vane platform, and a plurality of circumferentially spaced fasteners pass through apertures in the annular face plate into the support ring and are located in circumferential alignment with the cut-outs.

20. The turbine arrangement according to claim 17, characterized in that an inner seal member is affixed to the cylindrical flange and cooperates with the angel wing to define the second annular seal passage.