EP2535523A2 - Système d'étanchéité de turbine et procédé de son assemblage - Google Patents

Système d'étanchéité de turbine et procédé de son assemblage Download PDF

Info

Publication number
EP2535523A2
EP2535523A2 EP12171671A EP12171671A EP2535523A2 EP 2535523 A2 EP2535523 A2 EP 2535523A2 EP 12171671 A EP12171671 A EP 12171671A EP 12171671 A EP12171671 A EP 12171671A EP 2535523 A2 EP2535523 A2 EP 2535523A2
Authority
EP
European Patent Office
Prior art keywords
interstage seal
interstage
seal
wheel
turbine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP12171671A
Other languages
German (de)
English (en)
Inventor
Matthew Troy Hafner
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
Original Assignee
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Publication of EP2535523A2 publication Critical patent/EP2535523A2/fr
Withdrawn legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods

Definitions

  • the subject matter disclosed herein relates to gas turbines, and more specifically, to seals within turbines.
  • gas turbine engines combust a mixture of compressed air and fuel to produce hot combustion gases.
  • the combustion gases may flow through one or more turbine stages to generate power for a load and/or compressor.
  • a pressure drop may occur between stages, which may allow leakage flow of a fluid, such as combustion gases, through unintended paths.
  • Seals may be disposed between the stages to reduce fluid leakage between stages.
  • the seals may be subject to stresses, such as thermal stresses, which may bias the seals in axial and/or radial directions thereby reducing effectiveness of the seals. For example, seal deflection may increase the possibility of a rub condition between stationary and rotating components.
  • the present invention resides in a system including a multi-stage turbine.
  • the multi-stage turbine includes a first turbine stage including a first wheel having a plurality of first blade segments spaced circumferentially about the first wheel, a second turbine stage including a second wheel having a plurality of second blade segments spaced circumferentially about the second wheel, and an interstage seal extending axially between the first and second turbine stages.
  • the interstage seal is configured to be installed or removed while the first and second wheels remain in place in the respective first and second turbine stages.
  • the invention resides in a method including positioning a first recessed portion of an interstage seal about a first wheel rim of a turbomachine, pivoting a second recessed portion of the interstage seal toward an axial axis of the turbomachine, and moving the interstage seal along the axial axis toward a second wheel rim of the turbomachine to position the second recessed portion about the second wheel rim.
  • the present disclosure is directed to gas turbine engines that include interstage seals, wherein each interstage seal includes features to seal an interstage gap without the use of additional components, such as spacer wheels.
  • gas turbine engines that include such interstage seals may be less costly than engines using spacer wheels.
  • the gas turbine engine may include a first turbine stage that includes a first wheel that has a plurality of first blade segments spaced circumferentially about the first wheel, and a second turbine stage that includes a second wheel having a plurality of second blade segments spaced circumferentially about the second wheel.
  • the interstage seal may extend axially between the first and second turbine stages to seal the interstage gap.
  • embodiments of the interstage seal may be installed and removed without disassembling a rotor of the gas turbine engine.
  • the interstage seal may be configured to be installed or removed while the first and second wheels remain in place in the respective first and second turbine stages.
  • the interstage seal may include an inclined support rib that is configured to enable the interstage seal to pivot toward and away from an axial axis of the gas turbine engine without removal of the first wheel or the second wheel.
  • pivoting of the interstage seal may enable the interstage seal to be replaced without disturbing the rotor assembly.
  • a recessed portion of the interstage seal may be configured to enable the pivoting of the interstage seal.
  • FIG. 1 is a block diagram of an exemplary system 10 including a gas turbine engine 12 that may employ interstage seals configured to be installed or removed without rotor disassembly, as described in detail below.
  • the system 10 may include an aircraft, a watercraft, a locomotive, a power generation system, or combinations thereof.
  • the illustrated gas turbine engine 12 includes an air intake section 16, a compressor 18, a combustor section 20, a turbine 22, and an exhaust section 24.
  • the turbine 22 is coupled to the compressor 18 via a shaft 26.
  • air may enter the gas turbine engine 12 through the intake section 16 and flow into the compressor 18, which compresses the air prior to entry into the combustor section 20.
  • the illustrated combustor section 20 includes a combustor housing 28 disposed concentrically or annularly about the shaft 26 between the compressor 18 and the turbine 22.
  • the compressed air from the compressor 18 enters combustors 30, where the compressed air may mix and combust with fuel within the combustors 30 to drive the turbine 22.
  • the hot combustion gases flow through the turbine 22, driving the compressor 18 via the shaft 26.
  • the combustion gases may apply motive forces to turbine rotor blades within the turbine 22 to rotate the shaft 26.
  • the hot combustion gases may exit the gas turbine engine 12 through the exhaust section 24.
  • the turbine 22 may include a plurality of interstage seals, which may be installed or removed while rotating components of the turbine 22, such as wheels, remain in place. Thus, maintenance affecting the interstage seals may be performed without complete disassembly of the turbine 22.
  • FIG. 2 is a cross-sectional side view of an embodiment of the gas turbine engine 12 of FIG. 1 taken along the longitudinal axis 32.
  • the gas turbine 22 includes three separate stages 34.
  • Each stage 34 includes a set of blades 36 coupled to a rotor wheel 38 that may be rotatably attached to the shaft 26 ( FIG. 1 ).
  • the blades 36 extend radially outward from the rotor wheels 38 and are partially disposed within the path of the hot combustion gases.
  • Seals 40 extend between and are supported by adjacent rotor wheels 38.
  • the seals 40 may include recessed portions that fit about adjacent wheels 38 for support. The recessed portions may be configured to enable the seals 40 to pivot toward and away from the longitudinal axis 32 during installation or removal.
  • the seals 40 may be installed or removed while the rotor wheels 38 remain in place in the gas turbine engine 12.
  • the seals 40 may provide for improved cooling of the stages 34.
  • the gas turbine 22 is illustrated as a three-stage turbine, the seals 40 described herein may be employed in any suitable type of turbine with any number of stages and shafts.
  • the seals 40 may be included in a single stage gas turbine, in a dual turbine system that includes a low-pressure turbine and a high-pressure turbine, or in a steam turbine.
  • the seals 40 described herein may also be employed in a rotary compressor, such as the compressor 18 illustrated in FIG. 1 .
  • the seals 40 may be made from various high-temperature alloys, such as, but not limited to, nickel based alloys.
  • the compressed air from the compressor 18 is then directed into the combustor section 20 where the compressed air is mixed with fuel.
  • the mixture of compressed air and fuel is generally burned within the combustor section 20 to generate high-temperature, high-pressure combustion gases, which are used to generate torque within the turbine 22.
  • the combustion gases apply motive forces to the blades 36 to turn the wheels 38.
  • a pressure drop may occur at each stage 34 of the turbine 22, which may allow gas leakage flow through unintended paths.
  • the hot combustion gases may leak into the interstage volume between turbine wheels 38, which may place thermal stresses on the turbine components.
  • the interstage volume may be cooled by discharge air bled from the compressor or provided by another source. However, flow of hot combustion gases into the interstage volume may abate the cooling effects. Accordingly, the seals 40 may be disposed between adjacent wheels 38 to seal and enclose the interstage volume from the hot combustion gases. In addition, the seals 40 may be configured to direct a cooling fluid to the interstage volume or from the interstage volume toward the blades 36.
  • FIG. 3 is a cross-sectional side view of an embodiment of a pair of adjacent rotor stages 34 shown in FIG. 2 .
  • Hot fluids such as hot combustion gases or steam, with a flow path 56 (illustrated generally by an arrow) enters at an upstream side 58 and exits at a downstream side 60.
  • a flow path 56 illustrated generally by an arrow
  • a first turbine stage 62 is shown near the upstream side 58 and a second turbine stage 64 is shown near the downstream side 60.
  • the first turbine stage 62 includes a first wheel 66 with a plurality of first blade segments 68 extending radially outward 52 from a first wheel post portion 70 of the first wheel 66.
  • the first wheel post portion 70 is disposed along the circumference of the first wheel 66 and includes slots 72 (e.g., axial dovetail slots) for retaining lower segments (e.g., axial dovetail tabs 73) of the first blade segments 68.
  • the second turbine stage 64 includes a second wheel 74 with a plurality of second blade segments 76 extending radially outward 52 from a second wheel post portion 78 of the second wheel 74.
  • the second wheel post portion 78 is disposed along the circumference of the second wheel 74 and includes slots 80 (e.g., axial dovetail slots) for retaining lower segments (e.g., axial dovetail tabs 81) of the plurality of second blade segments 76.
  • slots 80 e.g., axial dovetail slots
  • lower segments e.g., axial dovetail tabs 81
  • approximately 50 to 150 first and second blade segments 68 and 76 may be mounted and spaced circumferentially 54 around the first and second wheels 66 and 74 and a corresponding axis of rotation (extending generally in the direction indicated by arrow 50).
  • methods other than the slots and tabs described above may be used to couple the first and second blade segments 68 and 76 to the first and second wheels 66 and 74.
  • the interstage seal 40 extends between the first and second adjacent wheels 66 and 74 and is mechanically supported by the first and second turbine stages 62 and 64.
  • an annular interstage seal assembly 41 (as shown in FIG. 9 ) may include a plurality of interstage seal segments 40 disposed about the longitudinal axis 32 of the gas turbine engine 12.
  • the interstage seal assembly 41 is a segmented annular seal assembly.
  • the interstage seal segment 40 includes an outer bridge portion 82, or axial beam, disposed near the first and second pluralities of blade segments 68 and 76.
  • the outer bridge portion 82 is a structure that provides support for the interstage seal 40.
  • the interstage seal segment 40 also includes an inner bridge portion 84 disposed near the first and second wheels 66 and 74.
  • the inner bridge portion 84 also provides support for the inner stage seal 40.
  • the inner bridge portion 84 may have a catenary shape, e.g., a curved annular shape, configured to provide additional strength to the interstage seal 40.
  • the interstage seal 40 includes one or more intermediate supports 86, or radial beams, which provide support for the inner stage seal 40 in the radial direction 52. As illustrated in FIG. 3 , the intermediate supports 86 may be generally aligned with the radial direction 52.
  • the interstage seal 40 may include three, four, five, six, or more intermediate supports 86.
  • the intermediate supports 86 may be disc-shaped structures that generally taper in the radial direction 52. In other words, the intermediate supports 86 may be thicker near the inner bridge portion 84 than near the outer bridge portion 82.
  • the interstage seal 40 may also include an inclined support rib 88, or support beam, that may be disposed between the inner and outer bridge portions 82 and 84. As shown in FIG. 3 , the inclined support rib 88 may be inclined with respect to the radial direction 52.
  • the interstage seal 40 may also include an optional inclined support portion 90 to provide additional support for the portion of the outer bridge portion 82 not supported by the inclined support rib 88. In certain embodiments, the inclined support portion 90 may be omitted. In further embodiments, the inclined support portion 90 may have a generally triangular cross sectional shape.
  • the inclined support rib 88 may be inclined with respect to the radial direction 52, the inclined support rib 88 may form an outer angle 92 with the outer bridge portion 82 and an inner angle 94 with the inner bridge portion 84.
  • the outer and inner angles 92 and 94 may be acute angles of less than approximately 90 degrees.
  • the outer and inner angles 92 and 94 may be between approximately 10 to 80 degrees, 20 to 70 degrees, 30 to 60 degrees, or 40 to 50 degrees.
  • the outer and inner angles 92 and 94 may be less than approximately 75 degrees.
  • the inclined support rib 88 enables the interstage seal 40 to be pivoted during installation and removal from the gas turbine engine 10.
  • the interstage seal 40 may be installed or removed without removal of the first and second wheels 66 and 74.
  • the interstage support portion 90 may be coupled to the inner bridge portion 84 instead of the outer bridge portion 82.
  • Seal cavities 96 may be formed in the interstage seal 40 between the intermediate supports 86.
  • the seal cavities 96 may enable a cooling fluid, such as air, to circulate between the first and second turbine stages 62 and 64 as discussed in detail below.
  • Recessed portions 98 may be formed between the outer and inner bridge portions 82 and 84 near the ends of the interstage seal 40 facing toward the first and second turbine stages 62 and 64.
  • the intermediate supports 86 and the inclined support rib 88 may not be located at the ends of the outer and inner bridge portions 82 and 84.
  • the recessed portions 98 are formed in the spaces surrounded by the intermediate supports 86, the inclined support rib 88, and the outer and inner bridge portions 82 and 84.
  • the seal cavities 96 may have a variety of cross sectional shapes depending on the configuration of the intermediate supports 86 and the inclined support rib 88.
  • the seal cavities 96 may have rectangular, square, triangular, circular, oval, or other suitable cross sectional shapes.
  • the recessed portions 98 may have a variety of cross sectional shapes, such as, but not limited to, rectangular, square, triangular, circular, oval, or other suitable shapes.
  • the inclined support portion 90 may occupy part of the recessed portion 98 adjacent to the inclined support rib 88. In other embodiments, the inclined support portion 90 may be omitted.
  • the recessed portions 98 may at least partially fit over portions of the first and second turbine stages 62 and 64. In other words, portions of the first and second wheels 66 and 74 may extend into the recessed portions 98 to enable pivoted motion of the interstage seal 40 and/or enable installation and removal of interstage seal 40 without removal of the first and second wheels 66 and 74.
  • a labyrinth seal 100 may be disposed adjacent to the interstage seal 40 and between the first and second turbine stages 62 and 64.
  • the labyrinth seal 100 may be configured to help block axial leakage of the hot combustion gases 56.
  • the labyrinth seal 100 may include an abradable coating 102 on the surface facing toward the interstage seal 40.
  • the interstage seal 40 may include one or more teeth 104 disposed adjacent to the abradable coating 102. During operation of the gas turbine engine 10, the teeth 104 may be in close proximity to the abradable coating 102 to help block axial leakage of the hot combustion gases 56 between the first and second turbine stages 62 and 64.
  • the abradable coating 102 may be configured to partially abrade when in contact with the teeth 104 to help prevent damage to the teeth 104.
  • the abradable coating 102 may be softer than the teeth 104.
  • seals other than the labyrinth seal 100 may be used together with the interstage seal 40.
  • the portions of the outer bridge portion 82 that extends past the intermediate support 86 and the inclined support rib 88 may be referred to as end portions.
  • the outer bridge portion 82 may include a first end portion 106 and a second end portion 108.
  • the first and second end portions 106 and 108 may include optional centrifugal seals 110 to help block radial leakage of the hot combustion gases 56.
  • the first and second end portions 106 and 108 may include a recessed slot 111 to engage with the centrifugal seal 110.
  • the seal 110 may include a support rod 112, a curved support piece 114, and a seal rod 116. The support rod 112 of the centrifugal seal 110 may fit in the recessed slot 111.
  • the curved support piece 114 may be attached to the support rod 112.
  • the seal rod 116 may be attached to the end of the curved support piece 114.
  • centrifugal forces may cause the seal rod 116 to move away from the interstage seal 40 and toward the surfaces of the first and second turbine stages 62 and 64 facing the interstage seal 40.
  • the seal rod 116 may be in contact with the first and second blade segments 68 and 76 during operation of the gas turbine engine 10 to help block radial leakage of the hot combustion gases 56.
  • small gaps exist between the first and second end portions 106 and 108 of the interstage seal 40 and the first and second turbines stages 62 and 64.
  • the centrifugal seals 110 may be able to maintain contact with the first and second turbine stages 62 and 64 even during axial transients that may cause the gaps to increase or decrease during operation of the gas turbine engine 10.
  • the centrifugal seals 110 may be omitted or seals other than the centrifugal seals 110 may be used at the outer bridge portion 82 to provide for radial sealing.
  • the second end portion 108 may include a first support feature 118 configured to engage with a second support feature 120 disposed on one or more of the second blade segments 76.
  • the first support feature 118 may be a female alignment portion (e.g., a notch) and the second support feature 120 may be a male alignment portion (e.g., a tab).
  • the first support feature 118 may be the male alignment portion
  • the second support feature 120 may be the female alignment portion.
  • the first and second support features 118 and 120 may help to block radial movement of the interstage seal 40 in the direction 52 toward the axial axis 50 of the gas turbine engine 10 during installation or removal of the interstage seal 40.
  • first and second support features 118 and 120 may help to block circumferential movement of the interstage seal 40 in the direction 54 during operation of the gas turbine engine 10. Use of the first and second support features 118 and 120 during installation and removal of the interstage seal 40 is described in detail below.
  • the inner bridge portion 84 may also include end portions, specifically, a first end portion 124, and a second end portion 126.
  • the first end portion 124 may be configured to engage with a first wheel rim 128 of the first wheel 66 during operation of the gas turbine engine 10. Specifically, during operation of the gas turbine engine 10, centrifugal forces may move the interstage seal 40 in the radial direction 52 toward the first rim 128. Contact between the first end portion 124 and the first rim 128 may provide an additional seal against radial leakage of the hot combustion gases 56.
  • the first end portion 124 may include an axial stop 130 disposed in the recessed portion 98.
  • the axial stop 130 may be a structure configured to restrict movement of the interstage seal 40 in the axial direction 50 toward the first turbine stage 62.
  • the second end portion 126 may be configured to engage with a second wheel rim 132 of the second wheel 74 during operation of the gas turbine engine 10. Contact of the second end portion 126 and the second rim 132 may help block radial leakage of the hot combustion gases 56.
  • Lengths 125 and 127 of the first and second end portions 124 and 126 may be selected to provide sufficient crush stress and clearance for assembly and removal for the interstage seal 40 depending on the selected materials.
  • the lengths 125 and 127 may be between approximately 5 mm to 50 mm, 10 mm to 25 mm, or 15 mm to 20 mm. Each of the lengths 125 and 127 may be between approximately 5 percent to 40 percent, 10 percent to 25 percent, or 15 percent to 20 percent of an overall length 136 of the interstage seal 40.
  • the interstage seal assembly 41 of which the interstage seal 40 is one segment of the assembly 41, is annularly disposed (in the circumferential direction 54) between the first and second wheels 66 and 74.
  • the first and second wheels 66 and 74 form annular structures with the interstage seal assembly 41 extending as an annular structure between the first and second wheels 66 and 74.
  • the first and second wheels 66 and 74 and the interstage seal assembly 41 rotate about a common axis.
  • the interstage seal assembly 41 may include a 360-degree segmented (e.g., 2 to 100 segments) circular structure that attaches to adjacent first and second wheels 66 and 74 to form a wall that thermally isolates an interstage volume or wheel cavity 134 that forms an air-cooling chamber.
  • FIGS. 4-6 illustrated various steps that may be performed during installation of the interstage seal 40. Removal of the interstage seal 40 may be accomplished by performing these steps in reverse.
  • FIG. 4 illustrates a partial cross sectional side view of the interstage seal 40 being pivoted between the first and second turbine stages 62 and 64. As shown in FIG. 4 , the second blade segments 76 have been removed to facilitate the installation of the interstage seal 40.
  • the first blade segments 68 remain in place in the first stage 62.
  • the interstage seal 40 may be installed by removing the first blade segments 68, with the second blade segments 76 remaining in place in the second stage 64.
  • installation of the interstage seal 40 may not involve removal of both the first and second blade segments 68 and 76, thereby substantially simplifying the installation or removal of the interstage seal 40.
  • the first and second wheels 66 and 74 remain in place during installation (and removal) of the interstage seal 40, thereby substantially simplifying the installation or removal of the interstage seal 40.
  • the second end portion 126 is positioned, or hooked, under the second rim 132.
  • a corner 148 formed between the inclined support rib 88 and the inner bridge portion 84 is placed adjacent to, in an overlapping relationship with, the second wheel rim 132.
  • an overlap 149 of the corner 148 and the second rim 132 exists in the axial 50 and radial 52 directions.
  • the interstage seal 40 may be pivoted, or rotated, about the corner 148 in the direction of the arrow 150 toward the axial axis 50.
  • an outer edge 151 of the first end portion 124 may follow an arc 152 as the interstage seal 40 is moved in the direction 150.
  • the overlap 149 provides sufficient clearance to enable the outer edge 151 to clear the first wheel rim 128 as the interstage seal 40 pivots toward the axial axis 50.
  • the inclined support rib 88 may be substantially parallel to a face 153 of the second wheel post portion 78.
  • the configuration of the inclined support rib 88 and the recessed portion 90 enables the overlap 149 to be greater, thereby enabling the outer edge 151 to clear the first wheel rim 128 as indicated by the dashed line 152.
  • the inclined support portion 90 may overlap a portion 154 of the second wheel support post 78 when viewed cross-sectionally. As described in detail below, two or more inclined support portions 90 may surround the second wheel support post 78.
  • FIG. 5 is a partial cross-sectional side view of the interstage seal 40 being moved along the axial axis 50.
  • the interstage seal 40 has been rotated such that the first end portion 124 may be moved under the first wheel rim 128 as indicated by arrow 170.
  • the second end portion 126 may remain overlapping with the second wheel rim 132.
  • the corner 148 may be adjacent to the second wheel rim 132.
  • the overlap 149 enables the first end portion 124 to move under the first wheel rim 128 while the overlap 149 is maintained between the second end portion 126 and the second wheel rim 132.
  • the axial stop 130 may contact the first wheel rim 128 to block further axial movement 50 of the interstage seal 40.
  • the corner 148 moves at least partially away from the second wheel rim 132.
  • FIG. 6 is a partial cross-sectional side view of the interstage seal 40 illustrating the completion of the installation process.
  • the axial stop 130 is adjacent to the first wheel rim 128, and the corner 148 has moved away from the second wheel rim 132.
  • the first end portion 124 remains axially 50 overlapped with the first wheel rim 128 and the second end portion 126 remains axially 50 overlapped with the second wheel rim 132.
  • the second blade segment 76 may be moved in the direction of arrow 180 toward the interstage seal 40.
  • the second support feature 120 may be engaged with the first support feature 118.
  • the interstage seal 40 may be blocked from moving toward or away from the axial axis 50 of the gas turbine engine 10.
  • the interstage seal 40 may be self-supporting throughout the rest of the installation process without having to hold or restrain the interstage seal 40 from moving.
  • the labyrinth seal 100 may be moved in the direction of arrow 182 toward the interstage seal 40.
  • the labyrinth seal 100 may be coupled to the case of the gas turbine engine 10 and thus, the labyrinth seal 100 may be installed when the case of the gas turbine engine 10 is mounted.
  • the centrifugal seals 110 (if used) may move outward to radially seal the gaps between the interstage seal 40 and the first and second turbine stages 62 and 64.
  • FIG. 7 is a front perspective view of an embodiment of one interstage seal segment 40 of the interstage seal assembly 41.
  • the axial stop 130 includes a first face 200 and a second face 202.
  • the first face 200 may face toward the first turbine stage 62 and the second face 202 may face toward the interstage seal 40.
  • the first face 200 may be generally flat to correspond to the first wheel 66.
  • the second face 202 may be curved to provide a greater support for attachment of the axial stop 130 to the first end portion 124.
  • the second face 202 may be flat.
  • the axial stop 130 may have a width 204 that is approximately the same as a width of the interstage seal segment 40.
  • the axial stop 130 may be defined by a height 206, which may be selected to provide sufficient surface area for the axial stop 130 to help block undesired axial movement of the interstage seal 40.
  • FIG. 8 is a rear perspective view of an embodiment of one interstage seal segment 40 of the interstage seal assembly 41.
  • the interstage seal 40 includes two inclined support portions 90. Specifically, the two interstage support portions 90 are located on outer sides 218 of the interstage seal 40. Thus, an interstage support portion gap 220 exists between the two inclined support portions 90.
  • the second wheel support post 78 may fit into the interstage support portion gap 220.
  • the interstage seal 40 may be pivoted about the second wheel support post 78 during installation or removal of the interstage seal 40. By pivoting about the second wheel support post 78, the second stage 64 may remain in place during maintenance of the interstage seal 40.
  • each of the inclined support portions 90 is defined by a width 222. Together, the widths 222 of the inclined support portions 90 and the interstage support portion gap 220 may be approximately the same as the width 204 of the interstage seal 40.
  • the inclined support portions 90 may be defmed by a height 224. As shown in FIG. 8 , the height 224 is less than a height of the interstage seal 40. In other embodiments, the height 224 may be smaller or greater depending on the amount of incline of the inclined support rib 88 and the support desired for the outer bridge portion 82. For example, if the angle 94 is smaller, the inclined support portions 90 may have greater heights 224 to provide additional support for the outer bridge portion 82. In other embodiments, only one inclined support portion 90 may be provided near the center of the interstage seal 40. In further embodiments, the interstage seal 40 may include more than two inclined support portions 90. Some embodiments may omit the inclined support portion 90.
  • FIG. 9 is a front view of three adjacent interstage seals 40 of the segmented interstage seal assembly 41.
  • the assembly 41 may include a plurality of interstage seals 40, such as 2 to 100 seals 40, disposed adjacent to one another to form a complete 360-degree ring about the longitudinal axis 32 of the gas turbine engine 10. As shown in FIG. 9 , each of the interstage seals 40 is arcuate in the circumferential direction 54.
  • the assembly 41 of the interstage seals 40 may include outer seals 240 and inner seals 242.
  • Axial slots 246 may be formed in the outer and inner bridge portions 82 and 84 to accommodate the outer and inner axial seals 240 and 242.
  • outer and inner seals 240 and 242 extend in the axial direction 50 along the axial slots 246.
  • Outer and inner seals 240 and 242 may be installed between each of the interstage seals 40 of the interstage seal assembly 41 as discussed above.
  • the outer and inner axial seals 240 and 242 may help to block radial leakage of the hot combustible gases 56.
  • FIG. 10 is a side view of an embodiment of the interstage seal 40.
  • the outer and inner seals 240 and 242 are disposed in axial slots 246 that run along the outer and inner bridge portions 82 and 84.
  • the interstage seal 40 may include one or more cooling passages 260 to enable the cooling fluid to flow toward the first and second blade segments 68 and 76.
  • the cooling passages 260 may enable the cooling fluid to flow from the interstage volume 134 toward the first and second blade segments 68 and 76.
  • the cooling fluid may flow from a casing structure of the gas turbine engine 10 into the first and second blade segments 68 and 76 through the cooling passages 260.
  • the cooling passages 260 may be formed in the outer and inner bridge portions 82 and 84, the intermediate support beams 86, and/or the inclined support rib 88.
  • the cooling passages 260 formed in the outer and inner bridge portions 82 and 84 may enable the cooling fluid to flow from the interstage volume 134 or the casing structure to the first and second blade segments 68 and 76.
  • the cooling passages 260 formed in the intermediate support beams 86 and/or the inclined support rib 88 may enable the cooling fluid to flow between the recessed portions 98 and the seal cavities 96.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Sealing Using Fluids, Sealing Without Contact, And Removal Of Oil (AREA)
EP12171671A 2011-06-17 2012-06-12 Système d'étanchéité de turbine et procédé de son assemblage Withdrawn EP2535523A2 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/163,418 US20120321437A1 (en) 2011-06-17 2011-06-17 Turbine seal system

Publications (1)

Publication Number Publication Date
EP2535523A2 true EP2535523A2 (fr) 2012-12-19

Family

ID=46245938

Family Applications (1)

Application Number Title Priority Date Filing Date
EP12171671A Withdrawn EP2535523A2 (fr) 2011-06-17 2012-06-12 Système d'étanchéité de turbine et procédé de son assemblage

Country Status (3)

Country Link
US (1) US20120321437A1 (fr)
EP (1) EP2535523A2 (fr)
CN (1) CN102852566A (fr)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2957721A1 (fr) * 2014-06-19 2015-12-23 United Technologies Corporation Systèmes et procédés de distribution d'air de refroidissement dans des turbines à gaz
EP2935837A4 (fr) * 2012-12-19 2016-01-13 United Technologies Corp Joint d'étanchéité segmenté de turbine à gaz
EP3088662A1 (fr) * 2015-02-20 2016-11-02 General Electric Company Joint interétage de turbine à plusieurs étages et procédé d'assemblage
EP3068997A4 (fr) * 2013-11-11 2017-07-05 United Technologies Corporation Joint d'étanchéité segmenté pour moteur à turbine à gaz
EP3287595A1 (fr) 2016-08-25 2018-02-28 Siemens Aktiengesellschaft Rotor dote d'un anneau d'etancheite segmente
DE102016215983A1 (de) 2016-08-25 2018-03-01 Siemens Aktiengesellschaft Rotor mit geteiltem Dichtungsring
US9920652B2 (en) 2015-02-09 2018-03-20 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US10634055B2 (en) 2015-02-05 2020-04-28 United Technologies Corporation Gas turbine engine having section with thermally isolated area

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9080447B2 (en) * 2013-03-21 2015-07-14 General Electric Company Transition duct with divided upstream and downstream portions
EP2995778B1 (fr) * 2014-09-12 2020-10-28 United Technologies Corporation Procédé et ensemble permettant de réduire la chaleur secondaire dans un moteur à turbine à gaz
US10662793B2 (en) * 2014-12-01 2020-05-26 General Electric Company Turbine wheel cover-plate mounted gas turbine interstage seal
FR3047075B1 (fr) * 2016-01-27 2018-02-23 Safran Aircraft Engines Piece de revolution pour banc d'essai de turbine ou pour turbomachine, banc d'essais de turbines comprenant ladite piece, et procede les utilisant
EP3540180A1 (fr) * 2018-03-14 2019-09-18 General Electric Company Conduites de purge de cavité inter-étage
FR3096722B1 (fr) * 2019-05-29 2021-12-03 Safran Aircraft Engines Joint d’étanchéité dynamique pour turbomachine comprenant une pièce abradable multicouche
US11519286B2 (en) * 2021-02-04 2022-12-06 General Electric Company Sealing assembly and sealing member therefor with spline seal retention

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2691749B1 (fr) * 1992-05-27 1994-07-22 Snecma Dispositif d'etancheite entre des etages d'aubes et un tambour tournant notamment pour eviter les fuites autour des etages d'aubes de redresseur .
FR2825748B1 (fr) * 2001-06-07 2003-11-07 Snecma Moteurs Agencement de rotor de turbomachine a deux disques aubages separes par une entretoise
US6558118B1 (en) * 2001-11-01 2003-05-06 General Electric Company Bucket dovetail bridge member and method for eliminating thermal bowing of steam turbine rotors
US6761529B2 (en) * 2002-07-25 2004-07-13 Mitshubishi Heavy Industries, Ltd. Cooling structure of stationary blade, and gas turbine
GB2438858B (en) * 2006-06-07 2008-08-06 Rolls Royce Plc A sealing arrangement in a gas turbine engine
US20090238683A1 (en) * 2008-03-24 2009-09-24 United Technologies Corporation Vane with integral inner air seal

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10138751B2 (en) 2012-12-19 2018-11-27 United Technologies Corporation Segmented seal for a gas turbine engine
EP2935837A4 (fr) * 2012-12-19 2016-01-13 United Technologies Corp Joint d'étanchéité segmenté de turbine à gaz
EP3068997A4 (fr) * 2013-11-11 2017-07-05 United Technologies Corporation Joint d'étanchéité segmenté pour moteur à turbine à gaz
US10287905B2 (en) 2013-11-11 2019-05-14 United Technologies Corporation Segmented seal for gas turbine engine
EP2957721A1 (fr) * 2014-06-19 2015-12-23 United Technologies Corporation Systèmes et procédés de distribution d'air de refroidissement dans des turbines à gaz
US9915204B2 (en) 2014-06-19 2018-03-13 United Technologies Corporation Systems and methods for distributing cooling air in gas turbine engines
US10634055B2 (en) 2015-02-05 2020-04-28 United Technologies Corporation Gas turbine engine having section with thermally isolated area
US9920652B2 (en) 2015-02-09 2018-03-20 United Technologies Corporation Gas turbine engine having section with thermally isolated area
EP3088662A1 (fr) * 2015-02-20 2016-11-02 General Electric Company Joint interétage de turbine à plusieurs étages et procédé d'assemblage
US10337345B2 (en) 2015-02-20 2019-07-02 General Electric Company Bucket mounted multi-stage turbine interstage seal and method of assembly
EP3287595A1 (fr) 2016-08-25 2018-02-28 Siemens Aktiengesellschaft Rotor dote d'un anneau d'etancheite segmente
DE102016215983A1 (de) 2016-08-25 2018-03-01 Siemens Aktiengesellschaft Rotor mit geteiltem Dichtungsring
WO2018036818A1 (fr) 2016-08-25 2018-03-01 Siemens Aktiengesellschaft Rotor muni d'une bague d'étanchéité segmentée

Also Published As

Publication number Publication date
CN102852566A (zh) 2013-01-02
US20120321437A1 (en) 2012-12-20

Similar Documents

Publication Publication Date Title
EP2535523A2 (fr) Système d'étanchéité de turbine et procédé de son assemblage
EP2639409B1 (fr) Système d'étanchéité inter-étages de turbine
US8511976B2 (en) Turbine seal system
US10774668B2 (en) Intersage seal assembly for counter rotating turbine
US8419356B2 (en) Turbine seal assembly
US7207771B2 (en) Turbine shroud segment seal
EP2613000B1 (fr) Système de rétention axiale de segments rotatifs d'une turbine et procedé associé
EP2564032B1 (fr) Élément d'une turbine pourvu de joints lamellaires et procédé permettant de former un joint d'étanchéité contre les fuites entre une aube et un élément porteur
US9624784B2 (en) Turbine seal system and method
JP2015535565A (ja) タービンシュラウドの取り付け及び封止の構成
US10662793B2 (en) Turbine wheel cover-plate mounted gas turbine interstage seal
EP2650484A1 (fr) Configuration de joint inter-étages d'une turbine
US8235656B2 (en) Catenary turbine seal systems
EP2613010A1 (fr) Ensemble d'étanchéité à balais à double extrémité destiné à un compresseur
US20110163505A1 (en) Adverse Pressure Gradient Seal Mechanism
US8956120B2 (en) Non-continuous ring seal
CN103502577A (zh) 涡轮静叶片及燃气轮机
US9605553B2 (en) Turbine seal system and method
US10337345B2 (en) Bucket mounted multi-stage turbine interstage seal and method of assembly
JP6117612B2 (ja) 圧縮機及びガスタービン
JP2012013084A (ja) 回転機械を組み立てる方法及び装置

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A2

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

AX Request for extension of the european patent

Extension state: BA ME

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20150106