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.
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.
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.
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:
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
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 Ac to 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
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
The first annular cavity C1 is 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 C2 is 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 C2 is 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 C2 corresponds 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 C1 is in limited fluid communication with the annular flow path 18 via a first annular seal passage P1 between the first cylindrical seal wall 36 and the angel wing 48. In particular, the first annular seal passage P1 can be formed between a radially extending rim portion 50a 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 C1 is in limited fluid communication with the second annular cavity C2 via a second annular seal passage P2 between the angel wing 48 and the outer end 52 of the annular seal plate 44. In particular, the second annular seal passage P2 can 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 C1 and defines a support surface for the seal member 56.
An axial forward side 60 of the axial distal end 50a 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 34a that is in facing engagement with a planar facing surface 62a 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 64a that extend from the annular face plate 34 into the first annular cavity C1.
Referring to
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 C1 serves 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 C1 damps 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 (
The presence of the bolt heads 64a extending into the first annular cavity C1 further operates to decrease passage of the hot gases into the second annular cavity C2 in that the bolt heads 64a 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 64a as an effect of stationary bolt drag to rotor rotation. Windage can result in heating of the cooling air in the second annular cavity C2 such 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 68a, 68b, as illustrated in
Referring to
Referring to
Referring to
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.
Filing Document | Filing Date | Country | Kind |
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PCT/US2015/041056 | 7/20/2015 | WO | 00 |