Patent Publication Number: US-9404376-B2

Title: Sealing component for reducing secondary airflow in a turbine system

Description:
BACKGROUND OF THE INVENTION 
     The subject matter disclosed herein relates to turbine systems and, more particularly, to a sealing component for reducing secondary airflow in a turbine system. 
     Turbine components are typically directly exposed to high temperature gases, and therefore require cooling to meet their useful life. For example, some of the compressor discharge air is diverted from the combustion process for cooling rotor components of the turbine. Turbine buckets, blades and vanes typically include internal cooling channels therein which receive compressor discharge air or other cooling gases for cooling thereof during operation. In addition, turbine rotor disks which support the buckets are subject to significant thermal loads and thus also need to be cooled to increase their lifetimes. 
     The main flow path of the turbine is designed to confine combustion gases as they flow through the turbine. Turbine rotor structural components must be provided with cooling air independent of the main gas flow to prevent ingestion of the hot combustion gases therein during operation, and must be shielded from direct exposure to the hot flow path gas. Such confinement is accomplished by rotary seals positioned between the rotating turbine buckets to prevent ingestion or back flow of the hot air or gases into interior portions of the turbine rotor structure. Such rotary seals are insufficient to completely protect the interior components, such as the rotor structure, rotor and rotor disks, requiring the additional use of purge flows of cooling air into and through the rotor cavity. Such additional measures to protect the interior components increase the cost and complexity and hinder the performance of gas turbines. 
     BRIEF DESCRIPTION OF THE INVENTION 
     According to one aspect of the invention, a sealing component for reducing secondary airflow in a turbine system includes a first end segment configured to be disposed between, and retained in a radial direction by, a first land on a first rotor disk and a first turbine bucket platform operatively coupled to the first rotor disk. Also included is a second end segment configured to be disposed between, and retained in a radial direction by, a second land on a second rotor disk and a second turbine bucket platform operatively coupled to the second rotor disk. Further included is a main body portion extending axially from the first end segment to the second end segment. 
     According to another aspect of the invention, a gas turbine engine includes a compressor section and a combustor section. Also included is a turbine section having a first turbine bucket attached to a first rotor disk, a second turbine bucket attached to a second rotor disk, and a stationary turbine nozzle located axially between the first rotor disk and the second rotor disk. Further included is a sealing component extending axially between the first rotor disk and the second rotor disk. The sealing component includes a first end segment disposed between, and in contact with, a first axially extending land of the first rotor disk and a first platform of the first turbine bucket. The sealing component also includes a second end segment disposed between, and in contact with, a second axially extending land of the second rotor disk and a second platform of the second turbine bucket. The sealing component further includes a main body portion extending between the first end segment and the second end segment. 
     According to yet another aspect of the invention, a method of sealing a flow path of a gas turbine engine is provided. The method includes positioning a first end segment of a sealing component on a first axially extending land of a first rotor disk. The method also includes positioning a second end segment of the sealing component on a second axially extending land of a second rotor disk. The method further includes positioning a first platform of a first turbine bucket on the first end segment to radially retain the first end segment between the first axially extending land and the first platform. The method yet further includes positioning a second platform of a second turbine bucket on the second end segment to radially retain the second end segment between the second axially extending land and the second platform. 
     These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
         FIG. 1  is a schematic illustration of a gas turbine engine; 
         FIG. 2  is a side view illustration of a portion of a gas turbine engine including a sealing component; and 
         FIG. 3  is a flow diagram illustrating a method of sealing a flow path of the gas turbine engine. 
     
    
    
     The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings. 
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to  FIG. 1 , a turbine system, such as a gas turbine engine, for example, is schematically illustrated and generally referenced with numeral  10 . The gas turbine engine  10  includes a compressor section  12 , a combustor section  14 , a turbine section  16 , a rotor  18  and a fuel nozzle  20 . It is to be appreciated that one embodiment of the gas turbine engine  10  may include a plurality of compressors  12 , combustors  14 , turbines  16 , rotors  18  and fuel nozzles  20 . The compressor section  12  and the turbine section  16  are coupled by the rotor  18 . 
     The combustor section  14  uses a combustible liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run the gas turbine engine  10 . For example, fuel nozzles  20  are in fluid communication with an air supply and a fuel supply  22 . The fuel nozzles  20  create an air-fuel mixture, and discharge the air-fuel mixture into the combustor section  14 , thereby causing a combustion that creates a hot pressurized exhaust gas. The combustor section  14  directs the hot pressurized gas through a transition piece into a turbine nozzle (or “stage one nozzle”), and other stages of buckets and nozzles causing rotation of turbine blades within an outer casing  24  of the turbine section  16 . 
     Referring to  FIG. 2 , a portion of the turbine section  16  is illustrated in greater detail. The turbine section  16  includes alternating inter-stage nozzle stages  26  and turbine stages, such as a first turbine stage  28  and a second turbine stage  30 . A sealing component  32  is disposed between the first turbine stage  28  and the second turbine stage  30 . Although the embodiments described herein are described with reference to the turbine section  16  of the gas turbine engine  10 , the embodiments may also be utilized in conjunction with the compressor section  12  of the gas turbine engine  10 . 
     The first turbine stage  28  and the second turbine stage  30  each include respective rotor disks attached to a rotor shaft (not shown) that causes the rotor disks to rotate about a central axis. Specifically, the first turbine stage  28  includes a first rotor disk  34  and the second turbine stage includes a second rotor disk  36 . A plurality of blades or buckets is removably attached to an outer periphery of each rotor disk. For illustration purposes, a single turbine bucket for each stage is illustrated. In particular, a first turbine bucket  38  is attached to the first rotor disk  34  and a second turbine bucket  40  is attached to the second rotor disk  36 . The buckets are attached by any suitable mechanism, such as an axially extending dovetail connection. In one embodiment, the buckets each include a bucket platform configured to attach to the corresponding rotor disk. In the illustrated embodiment, the first turbine bucket  38  includes a first platform  42  and the second turbine bucket  40  includes a second platform  44 . As used herein, an “axial” direction is a direction parallel to the central axis, and a “radial” direction is a direction extending from the central axis and perpendicular to the central axis. An “outer” location refers to a location in the radial direction that is farther away from the central axis than an “inner” location. 
     The nozzle stage  26  includes a plurality of nozzle vanes  46  that are each operatively connected to the outer casing  24  of the turbine section  16 , such as a turbine shell or an outer support ring attached thereto, and extend radially toward the central axis. In one embodiment, each of the plurality of nozzle vanes  46  are attached to an inner support ring having a diameter less than a diameter of the outer support ring. 
     A sealing component  32  is included to reduce heated gas or air from leaking into interior portions of the turbine section  16  and away from a flow path  50  defined by the buckets and the nozzle stage. The sealing component  32  is disposed in a fixed position relative to the rotating rotor disks, and therefore rotates along with the rotor disks. As described in detail below, the sealing component  32  causes a sealing connection between the sealing component  32  and the buckets, such as the first turbine bucket  38  and the second turbine bucket  40 . 
     The sealing component  32  is typically a single, uniform structure shaped similar to a tied-arch bridge and configured to handle centrifugal forces associated with operation of the gas turbine engine  10 . Specifically, the sealing component  32  includes a main body portion  52  formed of a relatively planar portion  54 , an arched portion  56 , and a plurality of tie segments  58  connecting the relatively planar portion  54  and the arched portion  56 . The plurality of tie segments  58  forms at least one, but typically a plurality of hollow portions  60 . The plurality of hollow portions  60  reduces the overall weight and material cost of the sealing component  32 . 
     A first end segment  62  and a second end segment  64  are disposed at opposite axial ends of the sealing component  32 , such that the main body portion  52  extends axially from the first end segment  62  and the second end segment  64 . The first end segment  62  is disposed between the first turbine bucket  38  and a first land  68  of the first rotor disk  34 . As shown, the first land  68  extends axially in an aft direction. In particular, the first end segment  62  is “sandwiched” and thereby retained in a radial direction by portions of the first turbine bucket  38  and the first land  68 . In the illustrated embodiment, the first end segment  62  includes a first end  70  in contact with a radially outer face of the first land  34  and a second end  72  in contact with a radially inner face of the first platform  42 . Similarly, the second end segment  64  is “sandwiched” and thereby retained in a radial direction by portions of the second turbine bucket  40  and a second land  74  of the second rotor disk  36 . The second land  74  extends axially in a forward direction. The second end segment  64  includes a third end  76  in contact with a radially outer face of the second land  74  and a fourth end  78  in contact with a radially inner face of the second platform  44 . 
     The sealing component  32  extends between adjacent turbine bucket stages, such as between the first turbine stage  28  and the second turbine stage  30 , as illustrated, to seal a region extending between the adjacent stages. The fitted relationship between the stages retains the sealing component  32  in an axial direction. In one embodiment, additional axial retention is provided with a hook arrangement. In such an embodiment, a portion of the first end segment  62  and/or the second end segment  64  is engaged with a receiving feature of the first land  68 , the second land  74 , the first platform  42  and/or the second platform  44 . 
     The sealing component  32  is cast or otherwise made from high temperature materials capable of withstanding elevated temperatures such as 1500° F. or greater. Examples of such materials include nickel based superalloys such as those alloys used for flow path components. Additionally or alternatively, the sealing component  32  may be actively cooled. To facilitate replacement of the sealing component  32 , typically the sealing component  32  is formed as a circumferential segment extending around a portion of an axis of rotation of the gas turbine engine  10 . 
     As illustrated in the flow diagram of  FIG. 3 , and with reference to  FIGS. 1 and 2 , a method of sealing a flow path of a gas turbine engine  100  is also provided. The gas turbine engine  10  and the sealing component  32  have been previously described and specific structural components need not be described in further detail. The method of sealing a flow path of a gas turbine engine  100  includes positioning a first end segment of a sealing component on a first axially extending land of a first rotor disk  102 . The method also includes positioning a second end segment of the sealing component on a second axially extending land of a second rotor disk  104 . A first platform of a first turbine bucket is positioned on the first end segment to radially retain the first end segment between the first axially extending land and the first platform  106 . A second platform of a second turbine bucket is positioned on the second end segment to radially retain the second end segment between the second axially extending land and the second platform  108 . 
     The devices, systems and methods described herein provide numerous advantages over alternative systems. For example, the devices, systems and methods provide the technical effect of increasing efficiency and performance of the turbine by reducing the number of components and by reducing or eliminating or reducing the need for cooling gas flows. For example, the sealing component  32  alleviates the need for spacer wheels used often employed to support other sealing components and assemblies. Furthermore, the prevention of air flow leakage into interior cavities of the turbine reduces the level of cooling flow required, thus improving turbine efficiency and reducing cost. 
     While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.