• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The present invention relates to a turbine blade (80) for a gas turbine. The turbine blade (80) has a wing profile (82), a tip shroud (88) arranged radially on the outside of the wing profile (82) and a sealing rail (96) which is arranged radially on the outside of the tip shroud (88) and is located in one Extends substantially tangential direction from a first end to a second end of the tip shroud (88). The sealing rail (96) has a maximum axial thickness at a point which is arranged between the first end and the second end and is offset from the airfoil (82) in the substantially tangential direction. The invention further relates to a gas turbine and a method for balancing the tip shroud (88) of a turbine blade (80) of a gas turbine.
    公开号:CH709266B1
    申请号:CH00227/15
    申请日:2015-02-20
    公开日:2020-06-30
    发明作者:Aaron Kareff Spencer；Robert Piersall Matthew；Werner Danielle
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    Technical field
    The present invention relates generally to gas turbines and relates more specifically to a turbine blade and to a method for balancing a tip shroud of a turbine blade of a gas turbine.
    Background of the Invention
    In a gas turbine, hot combustion exhaust gases from one or more combustors can flow through an intermediate piece and along a hot gas path of a turbine. A number of turbine stages may typically be arranged in series along the hot gas path so that the combustion exhaust gases flow through the first stage nozzles and blades and then through the later stages of the turbine. In this way, the directors can direct the combustion exhaust gases to the respective blades, causing the blades to rotate and drive a load, such as an electrical generator or the like. The combustion exhaust gases can be enclosed by stationary peripheral linings that enclose the blades, which can also help direct the combustion exhaust gases along the hot gas path.
    [0003] Certain turbine blades can have a peak shroud that is arranged radially on the outside opposite an airfoil. During the operation of the turbine, the tip shroud can prevent the failure of a wing profile in the event of vibration cracks due to vibration loads. However, loads in a throat area between the wing profile and the tip shroud can be introduced due to centrifugal forces acting on the tip shroud. According to some configurations, the turbine blade may also have a sealing rail that is disposed radially outward from the tip shroud and extends in a tangential direction with respect to a central axis of rotation of the turbine. The sealing rail can extend generally radially into a groove which is arranged in the relevant stationary peripheral cladding. In this way, the sealing rail can control or avoid leakage of combustion exhaust gases between the tip shroud and the stationary peripheral cladding. In addition, the sealing rail can reduce the bending of the lace shroud, although the additional mass of the sealing rail can increase loads on the throat area.
    According to a known configuration, the sealing rail may extend in a tangential direction from a first end to a second end of the tip shroud and the sealing rail may have an axial thickness that is constant along the tangential direction. Although such a configuration can control leakage over the lace shroud and reduce bending of the lace shroud, the additional mass of the sealing bar, particularly at the ends of the lace shroud, can significantly increase the stresses on the throat area. Increased loads at high operating temperatures can lead to a high creep speed at the tip shroud, which can reduce the turbine blade part life. In addition, the increased loads at elevated temperatures may reduce the turbine blade fatigue life. Furthermore, such a sealing rail configuration can provide challenges in achieving peak shroud balancing and frequency tuning of the turbine blade, which can also reduce the turbine blade part life.
    [0005] There is therefore a need for a turbine blade with an improved sealing rail configuration. In particular, such a sealing rail configuration can be used to achieve peak shroud balancing and frequency tuning of the turbine blade and can be optimized in order to achieve adequate sealing rail balancing while also providing the necessary sealing rail mass in order to support the peak shroud and to maintain the desired frequency limits. In this way, such a sealing rail configuration can increase the part life of the turbine blade and therefore reduce the occurrence of expensive repairs and shutdown of the turbine.
    Brief description of the invention
    [0006] The present invention thus provides a turbine blade for a gas turbine. The turbine blade may have an airfoil, a tip shroud located radially outward from the airfoil, and a sealing rail disposed radially outward of the tip shroud that extends in a substantially tangential direction from a first end to a second end of the tip shroud. The sealing rail may have a maximum axial thickness at a location located between the first end and the second end and offset from the airfoil in the substantially tangential direction with respect to a central axis of rotation of the turbine blade (100).
    [0007] The present invention also provides a method for balancing a shroud of a turbine blade of a gas turbine. The method can have the step of providing a sealing rail which is arranged radially outside of the tip shroud and extends essentially in a tangential direction from a first end to a second end of the tip shroud. The method may also include the step of varying the axial thickness of the sealing rail so that a maximum axial thickness is at a location between the first end and the second end and offset from a wing profile of the turbine blade in the substantially tangential direction with respect to a central one Axis of rotation of the turbine blade (100) is arranged.
    The method may include that the step of varying the axial thickness of the sealing rail comprises balancing the shroud around the airfoil.
    The method may include that the step of varying the axial thickness of the sealing bar includes distributing the mass of the sealing bar to obtain the desired frequency limits.
    [0010] The present invention also provides a gas turbine. The gas turbine may have a compressor, a combustion chamber in communication with the compressor, and a turbine in communication with the combustion chamber. The turbine may have a number of turbine blades arranged in a circumferential arrangement. Each of the turbine blades may have an airfoil, a tip shroud located radially outward from the airfoil, and a sealing rail disposed radially outward of the tip shroud and extending in a substantially tangential direction from a first end to a second end of the tip shroud. The sealing rail can have a maximum axial thickness at a point which is arranged between the first end and the second end and offset from the airfoil in the substantially tangential direction with respect to a central axis of rotation of the turbine blade (100).
    It may be advantageous that the sealing rail has a first end and a second end, the location of the maximum axial thickness being arranged at a first distance from the first end of the sealing rail and a second distance from the second end of the sealing rail, and the first distance being different from the second distance.
    It may be advantageous that the sealing rail has a first axial thickness at a first end of the sealing rail and a second axial thickness at the second end of the sealing rail, and wherein the first axial thickness is the same as the second axial thickness.
    In any embodiment of the invention mentioned above, it may be advantageous for the sealing rail to have a first axial thickness at the first end of the sealing rail and a second axial thickness at the second end of the sealing rail, and wherein the first axial thickness is different from the second axial thickness.
    In any embodiment of the invention mentioned above, it may be advantageous that the axial thickness of the sealing rail increases variably from the first end of the sealing rail to the point with maximum axial thickness and / or that the axial thickness of the sealing rail from the second end of the Sealing rail increases variably to the point with maximum axial thickness.
    It may be advantageous that the sealing rail has one or more regions with constant axial thickness, which are arranged between the point with maximum axial thickness and the first end of the sealing rail or the second end of the sealing rail.
    It may be advantageous that the point is offset with a maximum axial thickness relative to a pressure side of the wing profile.
    It may be advantageous that the point with maximum axial thickness is offset from a suction side of the airfoil.
    It may be advantageous if the lace shroud is connected to the airfoil via a fillet section, and the point with maximum axial thickness is offset in relation to the fillet section in the substantially tangential direction.
    It can be advantageous if the point with maximum axial thickness is aligned radially with a portion of the throat area.
    It may be advantageous that the sealing rail has an upstream surface and a downstream surface, and the location with maximum axial thickness through a first radially extending edge of the upstream surface and a second radially extending edge of the downstream -Area is formed.
    These and other features and improvements of the present application and the resulting patent will become apparent to those of ordinary skill in the art upon reading the following detailed description, and with consideration of the several drawings and the appended claims.
    Brief description of the drawings
    Fig. 1 is a schematic representation of a gas turbine having a compressor, a combustor and a turbine.
    Fig. 2 is a schematic illustration of a portion of a turbine, such as may be used in the gas turbine of Fig. 1, showing a number of turbine stages.
    Fig. 3 is a front view of a known turbine blade as it can be used in the turbine of Fig. 2, wherein the turbine blade has an airfoil, a lace shroud and a sealing rail.
    Fig. 4 is a top plan view of a portion of the turbine blade of Fig. 3 showing the tip shroud, sealing rail, and airfoil (illustrated by hidden lines).
    5 is a frontal view of an embodiment of a turbine blade as may be described herein and as used in a turbine of FIG. 2, the turbine blade having an airfoil, a lace shroud, and a sealing rail.
    Fig. 6 is a top view of a portion of a turbine blade of Fig. 5 showing the tip shroud, sealing rail, and airfoil (illustrated by hidden lines).
    Detailed description of the invention
    Reference is now made to the drawings, in which like reference numerals refer to like elements throughout the different views, with FIG. 1 showing a schematic circuit diagram of a gas turbine 10 as may be used herein. The gas turbine 10 can have a compressor 15. The compressor 15 compresses an incoming stream of air 20. The compressor 15 delivers the compressed flow of air 20 to a combustion chamber 25. The combustion chamber 25 mixes the compressed flow of air 20 with a pressurized flow of a fuel 30 and ignites the mixture to produce a flow of combustion exhaust gases 35. Although only a single combustor 25 is illustrated, the gas turbine 10 may have any number of combustors 25. The flow of combustion exhaust gases 35 is in turn delivered to a turbine 40. The flow of the combustion exhaust gases 35 drives the turbine 40 to generate mechanical work. The mechanical work generated in the turbine 40 drives a compressor 15 via a shaft 45 and an external load 50, such as an electric generator or the like. Other configurations and other components can be used herein.
    [0029] The gas turbine 10 may use natural gas, various types of synthesis gas and / or other types of fuels. The gas turbine 10 may be any of a number of different gas turbines offered by the General Electric Company of Schenectady, New York, including, but not limited to, such as a Series 7 or Series 9 heavy duty gas turbine or the like. The gas turbine 10 can have different configurations and can use other types of components. Other types of gas turbines can also be used herein. Multiple gas turbines, other types of turbines, and other types of power generation devices can be used herein together. Although a gas turbine 10 is shown herein, the present invention may also be applicable to any type of turbomachine, such as a steam turbine.
    FIG. 2 shows a schematic circuit diagram of a section of the turbine 40, which has a number of stages 52, which are arranged in a hot gas path 54 of the gas turbine 10. A first stage 56 may include a number of first stage circumferentially spaced vanes 58 and a number of first stage circumferentially spaced blades 60 disposed about a central axis CA of turbine 40. The first stage 56 may also include a first stage fairing 62 that extends circumferentially and surrounds the first stage blades 60. The first stage fairing 62 may have a number of fairing segments arranged adjacent to one another in an annular arrangement. Similarly, a second stage 64 may include a number of second stage nozzles 66, a number of second stage blades 68, and a second stage shroud 70 surrounding the second stage blades 68. In addition, a third stage 72 may include a number of third stage nozzles 74, a number of third stage blades 76, and a third stage cowl 78 that encloses the third stage blades 76. Although the portion of the turbine 40 is illustrated as including three stages 52, the turbine 40 may have any number of stages 52 arranged along the central axis CA of the turbine 40.
    3 and 4 show a turbine blade 80, as can be used in one of the stages 52 of the turbine 40. E.g. The blades 80 may be used in the second stage 64 or a later stage of the turbine 40. In general terms, the turbine blades 80 may have an airfoil 82, a shaft 84, and a platform 86 disposed between the airfoil 82 and the shaft 84. As described above, a number of blades 80 may be arranged in a circumferential arrangement within the stage 52 of the turbine 40. In this manner, the airfoil 82 of each blade 80 can extend radially with respect to the central axis CA of the turbine 40, while the platform 86 of each blade 80 extends tangentially with respect to the central axis CA of the turbine 40.
    As illustrated, the airfoil 82 may extend outward in a radial direction R from the platform 86 to a tip shroud 88 that is disposed about a tip end 90 of the blade 80. In some embodiments, the lace shroud 88 may be connected to the wing profile 82 via a fillet area 92. The shaft 84 may extend radially inward from the platform 86 to a root end 94 of the blade 80 so that the platform 86 generally forms an interface between the airfoil 82 and the shaft 84. As illustrated, platform 86 may be configured to extend substantially parallel to the central axis CA of turbine 40 during its operation. The shaft 84 may be configured to form a root structure, such as a dovetail, configured to secure the blade 80 to a turbine disk of the turbine 40. During operation of the turbine 40, the flow of combustion exhaust gases 35 moves along the hot gas path 54 and over the platforms 86 of the rotor blades 80, which together with the outer circumference of the turbine disk can essentially form the radially inner boundary of the hot gas path 54. In the same way, the tip shrouds 88 of the blades 80 can essentially form the radially outer boundary of the hot gas path 54.
    As illustrated in FIGS. 3 and 4, the rotor blade 80 can also have a sealing rail 96 which is arranged radially on the outside of the tip shroud 88. In some embodiments, sealing rail 96 may be formed integrally with tip shroud 88. As illustrated, the sealing rail 96 may extend in a tangential direction T with respect to the central axis CA of the turbine 40. In particular, the sealing rail 96 can extend in the tangential direction T from a first end 97 of the lace shroud 88 to a second end 98 of the lace shroud 88. As illustrated, the tip shroud 96 can have an axial thickness AT, measured in an axial direction A, which is constant along a length L of the sealing rail 96, measured in the tangential direction T. As described above, the sealing rail 96 can be substantially radial extend into a groove which is formed in an associated stationary peripheral cladding. In this way, the sealing rail 96 can control or prevent combustion exhaust gas leakage between the tip shroud 88 and the stationary peripheral cover during operation of the turbine 40. In addition, the sealing rail 96 can reduce the deflection of the tip shroud 88 during operation of the turbine 40, although the additional mass of the sealing rail 96 can increase the loads on the throat area 92.
    5 and 6 show an embodiment of a turbine blade 100 as described herein. The turbine blade 100 may be used in one of the stages 52 of the turbine 40 and may generally be configured in a manner that corresponds to the turbine blade 80 described above, although certain differences in structure and function are described hereinafter. E.g. The blade 100 can be used in the second stage 64 or in a later stage of the turbine 40. As shown, the blade 100 may include an airfoil 102, a shaft 104, and a platform 106 disposed between the airfoil 102 and the shaft. A number of blades 100 may be arranged in a circumferential arrangement within the stage 52 of the turbine 40. In this manner, the airfoil 102 of each blade may extend radially with respect to the central axis CA of the turbine 40, while the platform 106 of each blade 100 extends tangentially with respect to the central axis CA of the turbine 40.
    As illustrated, the airfoil 102 may extend outward in a radial direction R from the platform 106 to the tip shroud 108 disposed about a tip end 110 of the blade 100. In some embodiments, the lace shroud 108 may be connected to the wing profile 102 via a fillet area 112. The shaft 104 can extend radially inward from the platform 106 to a root end 114 of the moving blade 100, so that the platform 106 generally forms an interface between the airfoil 102 and the shaft 104. As illustrated, platform 106 may be configured to extend substantially parallel to the central axis CA of turbine 40 during its operation. The shaft 104 may be configured to form a root structure, such as a dovetail, that is configured to secure the rotor blade 80 to a turbine disk of the turbine 40. During operation of the turbine 40, the flow of the combustion exhaust gases 35 moves along the hot gas path 54 and over the platforms 106 of the rotor blades 100, which together with an outer circumference of the turbine disk form the radially inner boundary of the hot gas path 54. In the same way, the tip shrouds 108 of the blades 100 can essentially form the radially outer boundary of the hot gas path 54.
    As illustrated in FIGS. 5 and 6, the rotor blade 100 also has a sealing rail 116 which is arranged radially on the outside of the tip shroud 108. In some embodiments, sealing rail 116 may be integrally formed with tip shroud 108. As illustrated, the sealing rail 116 extends in a tangential direction T with respect to the central axis CA of the turbine 40. In particular, the sealing rail 116 extends in the tangential direction T from a first end 118 of the tip shroud 108 to a second end 120 of the tip shroud 108 In this way, the sealing rail 116 can have a length L, measured in the tangential direction T, which extends from a first end 122 of the sealing rail 116 to a second end 124 of the sealing rail. The first end 122 may be a front end and the second end 124 may be a rear end with respect to a direction of rotation of the sealing rail 116 about the central axis CA of the turbine. As illustrated, the sealing rail 116 can have an axial thickness, measured in an axial direction A, which varies along the length L of the sealing rail 116. The sealing rail 116 may have a first axial thickness AT1 at the first end 122 and a second axial thickness AT2 at the second end 124. In some embodiments, the first axial thickness AT1 may be the same as the second axial thickness AT2. In some embodiments, the first axial thickness AT1 may be greater than the second axial thickness AT2. In still other embodiments, the first axial thickness AT1 may be less than the second axial thickness AT2.
    According to some exemplary embodiments, the sealing rail 116 can have a pre-inclined configuration and therefore run with a small angle of inclination in relation to the tangential direction T, as shown in FIG. 6. In other words, the sealing rail 116 may extend in a substantially tangential direction with respect to the central axis CA of the turbine 40 as a result of the angle of inclination.
    As shown in Fig. 6, the sealing rail 116 may have a third axial thickness AT3 at a location between the first end 122 and the second end 124 of the sealing rail 116, and the third axial thickness AT3 may be greater than the first axial thickness AT1 and the second axial thickness is AT2. In particular, the third axial thickness AT3 can have a maximum axial thickness ATMAX of the sealing rail 116 at a location 126 with a maximum axial thickness which is arranged between the first end 122 and the second end 124. As shown, the maximum axial thickness location 126 may be located a first distance D1 from the first end 122 and a second distance D2 from the second end 124, the first distance Di being different from the second distance D2. In other words, the location 126 with maximum axial thickness can be arranged offset from the center point of the length L of the sealing rail 116. In some exemplary embodiments, the first distance D1 can be greater than the second distance D2. In other exemplary embodiments, the first distance D1 can be smaller than the second distance D2.
    According to some exemplary embodiments, as shown in FIG. 6, the location 126 with maximum axial thickness in the tangential direction T can be arranged offset from the airfoil 102 (illustrated by hidden lines). In particular, location 126 with maximum axial thickness can have a tangential offset TO, measured in tangential direction T, with respect to wing profile 102. In some embodiments, the location 126 with maximum axial thickness can be offset to a pressure side 128 of the airfoil 102, as shown. In other exemplary embodiments, location 126 with maximum axial thickness can be offset to a suction side 130 of wing profile 102. In some embodiments, the maximum axial thickness location 126 may be aligned with a portion of the throat portion 112 (illustrated by hidden lines). In other exemplary embodiments, the location 126 with maximum axial thickness can be offset from the throat section 112 in the tangential direction T.
    The axial thickness of the sealing rail 116 can increase from a first end 122 to the point 126 with maximum axial thickness and can also increase from the second end 124 to the point 126 with maximum axial thickness. In some embodiments, the thickness of the sealing bar may increase constantly from the first end 122 to the maximum axial thickness location 126 and may also increase constantly from the second end 124 to the maximum axial thickness location 126, as shown. In other exemplary embodiments, the sealing rail 116 can have one or more regions with a constant axial thickness which are arranged between the first end 122 and the point 126 with a maximum axial thickness and / or between the second end 124 and the point 126 with a maximum axial thickness. In still other embodiments, the axial thickness of the sealing rail 116 may vary from first end 122 to location 126 with maximum axial thickness and may also vary from second end 124 to location 126 with maximum axial thickness, as shown.
    [0041] The sealing rail 116 may have an upstream surface 132 and a downstream surface 134, each of which extends from the first end 122 to the second end 124. In accordance with certain exemplary embodiments, location 126 with maximum axial thickness can be formed by a first radially extending edge 136 of upstream surface 132 and a second radially extending edge 138 of downstream surface 134. In some embodiments, the first radially extending edge 136 and the second radially extending edge 138 may be aligned in the tangential direction T, as shown. In some embodiments, the location 126 of maximum axial thickness may be formed by a first flat surface of the upstream surface 132 and a second flat surface of the downstream surface 134. In some embodiments, the first flat surface and the second flat surface may be aligned in the tangential direction.
    The specific dimensions of the sealing rail 116, including the length L, the first axial thickness AT1, the second axial thickness AT2, the maximum axial thickness ATMAX, the first distance D1, the second distance D2 and the tangential offset TO, can be selected, in order to optimally achieve adequate balancing of the tip shroud 108 while also providing the necessary mass for supporting the tip shroud 108 and maintaining the desired frequency limits. Due to the changing axial thickness along the length L of the sealing rail 116, the mass of the sealing rail 116 close to the first and second ends 122, 124 can be selected such that only a necessary measure is provided to support the tip shroud 108 and the deflection is therefore reduced , without unnecessarily increased loads on the throat section 112. The maximum axial thickness ATMAX and the location 126 with maximum axial thickness, which is determined by the first distance D1, the second distance D2 and the tangential offset TO, can be selected in order to balance the To achieve peak shroud and provide the necessary mass for frequency tuning of the blade 100. Accordingly, the configuration of the sealing bar 116 can provide an improved optimization of the balancing and frequency adjustment of the tip shroud compared to the sealing bar 96 with a constant axial thickness.
    The exemplary embodiments described herein therefore provide a turbine blade with an improved sealing rail configuration to achieve tip shroud balancing and frequency tuning of the turbine blade. As described above, the sealing bar configuration can be optimized to achieve adequate tip shroud balancing while also providing the necessary sealing bar mass to support the tip shroud and maintaining the desired frequency limits. In this way, the sealing rail configuration can ultimately increase the part life of the turbine blade and therefore reduce the occurrence of expensive repairs and shutdown of the turbine.
    It should be understood that the foregoing relates only to certain embodiments of the present application and the resulting patent. Numerous changes and modifications can be made herein by one of ordinary skill in the art without departing from the general idea and scope of the invention as defined by the following claims and their equivalents.
    Reference list
    10 gas turbine 15 compressor 20 flow of air 25 combustion chamber 30 flow of fuel 35 flow of combustion gases 40 turbine 45 shaft 50 external load 52 turbine stages 54 hot gas path 56 first stage 58 first stage diffuser 60 first stage blades 62 first stage fairing 64 second stage 66 second stage nozzle 68 second stage blades 70 second stage fairing 72 third stage 74 third stage nozzle 76 third stage blades 78 third stage fairing 80 rotor blades 82 wing profile 84 shaft 86 platform 88 lace shroud 90 tip end 92 throat section 94 root end 96 sealing rail 97 first end 98 second end 100 turbine blade 102 wing profile 104 shaft 106 platform 108 tip shroud 110 spigot ends 112 throat section 114 root end 116 sealing rail 118 first end 120 second end 122 first end 124 second end 126 location with maximum axial thickness 128 pressure side 130 suction side 132 upstream surface 134 downstream surface 136 radially extending first edge 138 radially extending second edge
    权利要求:
    Claims (9)
    [1]
    1. turbine blade (100) for a gas turbine (10), the turbine blade (100) comprising:a wing profile (82, 102);a tip shroud (88, 108) arranged radially outside of the wing profile (82, 102); anda sealing rail (96, 116) which is arranged radially outside of the tip shroud (88, 108) and which extends in a substantially tangential direction from a first end (97, 118) to a second end (98, 120) of the tip shroud (88 , 108), the sealing rail (96, 116) having a maximum axial thickness at a point (126) which is between the first end (97, 118) and the second end (98, 120) of the tip shroud (88, 108 ) is arranged in a substantially tangential direction with respect to a central axis of rotation of the turbine blade (100).
    [2]
    2. The turbine blade (100) according to claim 1, wherein the sealing rail (96, 116) has a first end (122) and a second end (124), the location (126) having maximum axial thickness at a first distance from the first End (122) of the sealing rail (96, 116) and at a second distance from the second end (124) of the sealing rail (96, 116), and wherein the first distance is different from the second distance.
    [3]
    3. Turbine blade (100) according to claim 1 or 2, wherein the sealing rail (96, 116) at the first end (122) of the sealing rail (96, 116) a first axial thickness and at the second end (124) of the sealing rail (96 , 116) has a second axial thickness and the first axial thickness is the same as the second axial thickness.
    [4]
    4. Turbine blade (100) according to one of claims 1 or 2, wherein the sealing rail (96, 116) at the first end (122) of the sealing rail (96, 116) a first axial thickness and at the second end (124) of the sealing rail (96, 116) has a second axial thickness, and wherein the first axial thickness is different from the second axial thickness.
    [5]
    5. turbine blade (100) according to any one of the preceding claims, wherein the axial thickness of the sealing rail (96, 116) from the first end (122) of the sealing rail (96, 116) to the point (126) with maximum axial thickness increases unevenly and wherein the axial thickness of the sealing rail (96, 116) increases unevenly from the second end (124) of the sealing rail (96, 116) to the point (126) with maximum axial thickness.
    [6]
    6. Turbine blade (100) according to one of the preceding claims, wherein the sealing rail (96, 116) has one or more sections with constant axial thickness, which between the point (126) with maximum axial thickness and the first end (122) of the sealing rail (96, 116) or the second end (124) of the sealing rail (96, 116) are arranged.
    [7]
    7. Turbine blade (100) according to one of the preceding claims, wherein the sealing rail (96, 116) has an upstream surface (132) and a downstream surface (134), and wherein the location (126) with a maximum axial thickness by a radially extending first edge (136) of the upstream surface (132) and a radially extending second edge (138) of the downstream surface (134) is formed.
    [8]
    8. A method for balancing a tip shroud (88, 108) of a turbine blade (100) of a gas turbine (10), the method comprising:Providing a sealing rail (96, 116) which is arranged radially outside of the tip shroud (88, 108) and which extends in a substantially tangential direction from a first end (97, 118) to a second end (98, 120) of the tip shroud ( 88, 108) extends;Changing an axial thickness of the sealing rail (96, 116) such that a maximum axial thickness at a point (126) between the first end (97, 118) and the second end (98, 120) and in the substantially tangential direction with Reference to a central axis of rotation of the turbine blade (100) is arranged.
    [9]
    9. Gas turbine (10), comprising:a compressor (15);a combustion chamber (25) in communication with the compressor (15); anda turbine (40) in communication with the combustion chamber (25), the turbine (40) having a plurality of turbine blades (100) arranged in a circumferential arrangement, each of the turbine blades (100) comprising:a wing profile (82, 102);a lace shroud (88, 108) disposed radially outward of the airfoil (82, 102); anda sealing rail (96, 116) disposed radially outward of the tip shroud (88, 108) and extending in a substantially tangential direction from a first end to a second end of the tip shroud (88, 108), the sealing rail (96 , 116) has a maximum axial thickness at a point (126) which is offset from the airfoil (82, 102) between the first end (97, 118) and the second end (98, 120) and in the substantially tangential direction.with respect to a central axis of rotation of the turbine blade (100) is arranged.
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    同族专利:
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    CN105041387B|2018-06-08|
    DE102015102396A1|2015-08-20|
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    US9464530B2|2016-10-11|
    US20150233258A1|2015-08-20|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2019-05-31| NV| New agent|Representative=s name: FREIGUTPARTNERS IP LAW FIRM DR. ROLF DITTMANN, CH |
    优先权:
    申请号 | 申请日 | 专利标题
    US14/185,161|US9464530B2|2014-02-20|2014-02-20|Turbine bucket and method for balancing a tip shroud of a turbine bucket|
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