EP3653844A1 - Streifendichtung, ringsegment und verfahren für eine gasturbine - Google Patents

Streifendichtung, ringsegment und verfahren für eine gasturbine Download PDF

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Publication number
EP3653844A1
EP3653844A1 EP18206450.1A EP18206450A EP3653844A1 EP 3653844 A1 EP3653844 A1 EP 3653844A1 EP 18206450 A EP18206450 A EP 18206450A EP 3653844 A1 EP3653844 A1 EP 3653844A1
Authority
EP
European Patent Office
Prior art keywords
annular segment
strip seal
locking member
main body
cavity
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Withdrawn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
EP18206450.1A
Other languages
English (en)
French (fr)
Inventor
David Overton
Peter Welburn
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.)
Siemens AG
Original Assignee
Siemens AG
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 Siemens AG filed Critical Siemens AG
Priority to EP18206450.1A priority Critical patent/EP3653844A1/de
Publication of EP3653844A1 publication Critical patent/EP3653844A1/de
Withdrawn legal-status Critical Current

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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
    • F01D11/005—Sealing means between non relatively rotating elements
    • F01D11/006—Sealing the gap between rotor blades or blades and rotor
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • 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
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F05D2240/57—Leaf seals
    • 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
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position
    • F05D2260/36—Retaining components in desired mutual position by a form fit connection, e.g. by interlocking
    • 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
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position
    • F05D2260/38—Retaining components in desired mutual position by a spring, i.e. spring loaded or biased towards a certain position

Definitions

  • the present disclosure relates to a strip seal and an annular segment for a gas turbine.
  • the present disclosure further relates to a method of assembling a gas turbine using the strip seal and the annular segment.
  • Gas turbine engines which are a specific example of turbomachines, generally include a rotor with a number of rows of rotor blades which are fixed to a rotor shaft via rotor discs and rows of stator vanes therebetween, which are fixed to the casing of the gas turbine.
  • a hot and pressurised working fluid flows through the rows of rotor blades and stator vanes in the main passage of a gas turbine, it transfers momentum to the rotor and thus imparts a rotary motion to the rotor while expanding and cooling. Maintaining the working fluid within the main passage is therefore crucial for engine performance as well as component life.
  • the rotor disc assemblies and/or stator vanes are assembled from a plurality of individual components, i.e. annular segments.
  • assembling the individual components into an array causes gaps to be formed between adjacent segments.
  • gaps may be utilised by the working fluid to escape from the main passage, it is known to provide seals, for example in the form of strip seals, to seal these gaps.
  • the arrayed segments define slots where they are interfaced. These slots are open-ended so that strip seals can be inserted following assembly of the segments. However, the strip seals may slide out of the open end of the slot. Assembly of segments may therefore be difficult and time consuming. Moreover, it may not be possible to ensure all strip seals are seated correctly.
  • a strip seal (200) for use in a gas turbine comprising: an elongate main body (220) defining a longitudinal direction, the main body (220) comprising: a first end (221) and a second end (222) which delimit the main body (220) along the longitudinal direction, and a cavity (240) defined by the main body (220); and a locking member (260) which projects away from the main body (220); wherein the strip seal (260) is resiliently deformable to contain the locking member (260) in the cavity (240).
  • the strip seal (200) provides a sealing means which may be inserted into a slot formed by a pair of annular segments (300) and secured therein by means of the locking member (260) to maintain the strip seal (200) seated correctly.
  • the main body (220) may define: a central region (227) extending between the first end (221) and the second end (222), and a pair of equally-sized side regions (228, 229) flanking the central region (227); wherein the cavity (240) and the locking member (260) are located in one of the side regions (228, 229).
  • the central region (220) of the main body (100) is located in the gap between a pair of annular segments (300), the central region (220) seals the gap by virtue of its presence therein.
  • the central region (220) may better seal the gap.
  • the other one of the side regions (228, 229) may not be provided with a locking member (260).
  • a single locking member (260) may be sufficient for retaining the strip seal (200) in position. Moreover, the manufacturing of such a strip seal (200) and corresponding gas turbine annular segment (300) receiving the strip seal (200) may be improved.
  • the locking member (260) may be configured to project from a cavity edge (242) of the main body (220), and the main body (220) and the locking member (260) are provided at an oblique angle to one another.
  • the strip seal (200) is provided in a retroserrate configuration which defines an insertion direction.
  • the locking member (260) is caused to be depressed into the cavity (240). Assembly of the gas turbine may therefore be faster and more convenient.
  • the cavity (240) and the locking member (260) may be provided closer to the first end (221) than the second end (222) or closer to the second end (222) than the first end (221).
  • the cavity (240) may extend through the main body (220).
  • the strip seal (200) may be manufactured by suitable manufacturing processes, such as laser-cutting, of the cavity (240) into the strip seal (200) and deforming the strip seal (200) suitably so that the locking member (260) extends from the main body (220). This manufacturing process may be easier to carry out by cutting all the way through the strip seal (200) in order to form the cavity (240) and the locking member (260).
  • annular segment (300) for a turbine rotor assembly or stator vane assembly of a gas turbine comprising: a platform (320) having a leading end (321) and a trailing end (322) bounding the platform (320) along an axial direction; an aerofoil portion (340) extending from the platform (320) along a radial direction; a wedge face (360) provided between the leading end (321) and the trailing end (322); wherein the wedge face (360) defines: a linear track (362) extending along the wedge face (360), the linear track (362) having an open end (364); and a recess (366) formed along the linear track (362).
  • the annular segment (300) may be provided as a stator vane segment (300).
  • the linear track (362) may be formed in a radially outer platform (320) of the annular segment (300).
  • the strip seal (200) according to the present disclosure is applicable particularly to the outer platforms (320) of stator vane segment (300).
  • the annular segment (300) may be provided as a rotor blade (300).
  • the linear track (362) may be formed in a radially inner platform (320) of the rotor blade (300).
  • the strip seal (200) according to the present disclosure is also applicable to an assembly of blade-carrying rotor discs, particularly with reference to the radially inner platform (320).
  • annular segment assembly (400) for a gas turbine comprising: a strip seal (200) as set out above, and a first annular segment (300) and a second annular segment (300) as set out above; wherein a wedge face (360) of the first annular segment (300) and a wedge face (360) of the second annular segment (300) are interfaced so that: a gap (410) is defined between the wedge faces (360), a slot (430) is defined by a linear track (362) of the first annular segment (300) and a linear track (362) of the second annular segment (300); wherein the strip seal (200) is located in the slot (430), and the locking member (260) of the strip seal (200) is located in the recess (366) formed along the linear track (362) of the first annular segment (300) or the linear track (362) of the second annular segment (300).
  • a method of assembly of a gas turbine comprising: forming an annular segment assembly (400) by: aligning a first annular segment (300) and a second annular segment (300) as set out above, thereby interfacing a wedge face (360) of the first annular segment (300) and a wedge face (360) of the second annular segment (300); sealing the gap (410) between the wedge faces (360) of the first annular segment (300) and the second annular segment (300) by: providing a strip seal (200) as set out above; inserting the strip seal (200) into a slot (430) formed by the component assembly (400), thereby causing the strip seal (200) to resiliently deform to locate the locking member (260) within the cavity (240); displacing the strip seal (200) along the slot (430) until the locking member (260) of the strip seal (200) reaches the recess (366), thus causing the locking member (260) to extend into the recess (366).
  • the method of assembly may comprise: separating the first annular segment (300) and the second annular segment (300), and recovering the strip seal (200) in a substantially undamaged condition.
  • the strip seal (200) is recoverable in a re-usable condition and can therefore be used following the disassembly of annular segments (300). Moreover, this is applicable also to service and maintenance of a gas turbine.
  • strip seals may not be seated correctly and may therefore suffer damage during assembly and/or disassembly of the annular assembly.
  • Using the strip seal (200) according to the present disclosure allows the strip seals (100) to be recovered in a substantially undamaged condition.
  • the present disclosure relates to a strip seal for use in a turbomachine, such as a gas turbine, and a method of assembling a gas turbine provided with a strip seal.
  • Figure 1 shows an example of a gas turbine engine in a sectional view, which illustrates the nature of components according to the present disclosure (for example rotor blades) and the environment in which they operate.
  • the gas turbine engine comprises, in flow series, an inlet 62, a compressor section 64, a combustion section 66 and a turbine section 68, which are generally arranged in flow series and generally in the direction of a longitudinal or rotational axis 70.
  • the gas turbine engine further comprises a shaft 72 which is rotatable about the rotational axis 70 and which extends longitudinally through the gas turbine engine.
  • the rotational axis 70 is normally the rotational axis of an associated gas turbine engine. Hence any reference to “axial”, “radial” and “circumferential" directions are with respect to the rotational axis 70.
  • the shaft 72 drivingly connects the turbine section 68 to the compressor section 64.
  • air 74 which is taken in through the air inlet 62 is compressed by the compressor section 64 and delivered to the combustion section or burner section 66.
  • the burner section 66 comprises a burner plenum 76, one or more combustion chambers 78 defined by a double wall can 80 and at least one burner 82 fixed to each combustion chamber 78.
  • the combustion chambers 78 and the burners 82 are located inside the burner plenum 76.
  • the compressed air passing through the compressor section 64 enters a diffuser 84 and is discharged from the diffuser 84 into the burner plenum 76 from where a portion of the air enters the burner 82 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 86 or working gas from the combustion is channelled via a transition duct 88 to the turbine section 68.
  • the turbine section 68 may comprise a number of blade carrying discs 90 or turbine wheels attached to the shaft 72.
  • the turbine section 68 comprises two discs 90 which each carry an annular array of turbine assemblies 12, which each comprises an aerofoil 14 embodied as a turbine blade 100.
  • Turbine cascades 92 are disposed between the turbine blades 100.
  • Each turbine cascade 92 carries an annular array of turbine assemblies 12, which each comprises an aerofoil 14 in the form of guiding vanes (i.e. stator vanes 96), which are fixed to a stator 94 of the gas turbine engine.
  • Figure 2 shows an enlarged view of a stator vane 96 and rotor blade 100.
  • Arrows “A” indicate the direction of flow of combustion gas 86 past the aerofoils 96,100.
  • Arrows “B” show air flow passages provided for sealing, and arrows “C” indicate cooling air flow paths for passing through the stator vanes 96.
  • Cooling flow passages 101 may be provided in the rotor disc 90 which extend radially outwards to feed an air flow passage 103 in the rotor blade 100.
  • the combustion gas 86 from the combustion chamber 78 enters the turbine section 58 and drives the turbine blades 100 which in turn rotate the shaft 72 to drive the compressor.
  • the guiding vanes 96 serve to optimise the angle of the combustion or working gas 86 on to the turbine blades.
  • Figure 3 shows a view of the rotor blades 100 looking upstream, facing the flow "A" shown in Figure 2 .
  • Each rotor blade 100 comprises an aerofoil portion 104, a root portion 106 and a platform 108 from which the aerofoil extends.
  • the rotor blades 100 are fixed to the rotor disc 102 by means of their root portions 106, through which the flow passage 101 may extend.
  • the root portions 106 have a shape that corresponds to notches (or grooves) 109 in the rotor disc 90, and are configured to prevent the rotor blade 100 from detaching from the rotor disc 102 in a radial direction as the rotor disc 102 spins.
  • a strip seal 110 is provided to seal a region (or gap) between pairs of adjacent platforms 108.
  • the segment of the present disclosure may relate to a rotor blade or another component for a gas turbine engine, for example a nozzle guide vane.
  • Figures 4 and 5 illustrate a strip seal 200 according to the present disclosure.
  • Figure 4 is a perspective view of the strip seal 200
  • Figure 5 is a top down view of the strip seal 200.
  • the strip seal 200 may be used with a gas turbine engine, such as the exemplary gas turbine engine of Figure 1 .
  • the strip seal 200 is a sealing member configured to seal an elongate gap, completely or at least partially, by means of its presence therein. More particularly, the strip seal 200 may be inserted into the turbine assemblies 12, i.e. the rotor blades 100 and/or the stator vanes 96, to seal gaps.
  • the strip seal 200 comprises a main body 220.
  • the main body 220 has a generally elongate shape. In use, at least a portion of the main body 220 is physically located within a gap in order to seal said gap.
  • the main body 220 comprises a first end 221, a second end 222, and a pair of side edges 223, 224 extending between the first end 221 and the second end 222 of the main body 220.
  • the first end 221 and the second end 222 bound the main body 220 along a first direction, e.g. longitudinal direction.
  • the side edges 223, 224 bound the main body 220 along a second direction, e.g. lateral direction.
  • the main body 220 further comprises a first surface 225 (or 'top face') and a second surface 226 (or bottom face').
  • the first surface 225 and the second surface 106 delimit the main body 220 along a third direction, which may be perpendicular to both the longitudinal direction and the lateral direction.
  • the main body 220 has a substantially constant cross-section along its length, i.e. from the first end 221 to the second end 222.
  • the side edges 223, 224 are straight and arranged to be parallel.
  • the first surface 225 and the second surface 226 are flat and arranged to be parallel.
  • the main body 220 defines a central region 227 and a pair of side regions 228, 229, i.e. a 'centre' and two 'margins'. That is to say, the main body 220 may divided into geometric regions.
  • the central regions 227 and the side regions 228, 229 extend from first end 221 to the second end 222.
  • the side regions may be equally sized and arranged to flank the central region 227.
  • a cavity 240 is defined by the main body 220.
  • the cavity 240 extends through the main body 220, i.e. from the first surface 225 to the second surface 226.
  • the cavity 240 is a recess, for example a cut-out or notch, on one of the side edges 223, 224 of the main body 220.
  • the strip seal 200 comprises a locking member 260.
  • the locking member 260 is carried by the main body 220 and configured to project away from the main body 220. More particularly, the locking member 260 is configured to project beyond at least one of the side edges 223, 224 or the surfaces 225, 226. That is to say, the locking member 260 extends from the main body 220 along a non-longitudinal direction. In other words, the locking member 260 extends at an angle to the longitudinal direction. According to some examples, the locking member 260 projects, i.e. extends, from one of the side edges 223, 224, while according to other examples the locking member 260 projects from one of the surfaces 225, 226. According to the present example, the locking member 260 extends from the first surface 225. More particularly, the locking member 260 extends from an edge 242 of the cavity 240.
  • the strip seal 200 is resiliently deformable to contain the locking member 260 in the cavity 240. That is to say, the strip seal 200, or at least a portion of the strip seal 200, is physically deformable so that the locking member 260 may be located within the cavity 240. Additionally, the strip seal 200 is configured to resist the locking member 260 being located in the cavity 240. That is to say, the locking member 260 is biased to assume a configuration in which the locking member 260 is not contained in, and therefore extends out of and away from, the cavity 240. According to the present example, the locking member 260 is resiliently deformable to be located in the cavity 240. Put another way, the locking member 260 may be resiliently biased such that at rest it is at an angle to the main body 220 of the strip seal 200, but may be forced, i.e. deformed, to take a position within the cavity 240.
  • the cavity 240 and the locking member 260 are located in one of the side regions 228, 229. Therefore, the cavity 240 and the locking member 260 are located beside the central region 227 extending along the strip seal 200, as opposed to being located within the central region 227.
  • the cavity 240 and the locking member 260 are provided at one of the side edges 223, 224 of the main body 220.
  • the central region 227 and the other side region 229 are provided without a locking member 260 extending therefrom. That is to say, the central region 227 and the other side region 229 are flat (or 'continuous').
  • the cavity 240 and the locking member 260 are provided towards the first end 221 or the second end 222 of the strip seal 200. According to other examples, the cavity 240 and the locking member 260 may not be provided at the first end 221 and the second end 222 of the strip seal 200. That is to say, the cavity 240 and locking member 260 may only be provided towards one of the first end 221 or the second end 222.
  • the strip seal 200 is provided in a retroserrate configuration. That is to say, the locking member 260 is angled towards the first end 221 or the second end 222 of the strip seal 200. Accordingly, the locking member 260 is configured to be depressed into the cavity 240 by means of being inserted in the insertion direction, and to resist removal along the opposite direction.
  • Figures 6 and 7 illustrate an annular segment 300 according to the present disclosure.
  • Figure 6 shows a perspective view of the annular segment 300
  • Figure 7 shows a partial perspective view of the annular segment 300.
  • the annular segment 300 may be used with a gas turbine engine, such as the exemplary gas turbine engine of Figure 1 .
  • the annular segment 300 is provided as a turbine stator vane, for example a nozzle guide vane.
  • the annular segment may be provided as a rotor blade, for example as shown in Figures 2, 3 .
  • the structure of a rotor blade and stator vane are different, they comprise common features as described in the following.
  • the annular segment 300 comprises a platform 320.
  • the platform 320 has a leading end 321 which is spaced apart from a trailing end 322 along the axial direction.
  • the leading end 321 and the trailing end 322 bound the platform 320 along the axial direction.
  • the annular segment 300 further comprises an aerofoil portion 340 extending from the platform 320 along the radial direction.
  • the annular segment 300 further comprises a pair of wedge faces 360 bounding the annular segment 300 along the circumferential direction.
  • a wedge face 360 of the annular segment 300 is configured to be interfaced with a wedge face 360 of a corresponding annular segment 300 when arrayed to form an annular assembly, i.e. rotor blade or stator vane.
  • Each wedge face 360 defines a linear track 362 (or 'straight' track) extending along the wedge face 360.
  • the linear track 362 is configured to receive the strip seal 200. More particularly, the linear track 362 may receive a side edge 223, 224 of the strip seal 200.
  • the linear track 362 comprises an open end 364.
  • the annular segment 300 is configured to receive the strip seal 200 through the open end 364.
  • the open end 364 is accessible along the axial direction and/or the radial direction.
  • the open end 364 remains accessible when the annular segments 300 have been arrayed (i.e. assembled to form an array) to form an annular assembly a gas turbine, i.e. a stator vane assembly or a rotor blade assembly.
  • the linear track 362 is formed in radially inwards from the aerofoil portion 340, e.g. the root of the annular segment 300.
  • the linear track 362 is formed radially outwards from the aerofoil portion 340, in the outer platform of the nozzle guide vane.
  • At least one of the wedge faces 360 defines a recess 366 formed along the linear track 362.
  • both wedge faces 360 define a recess 366 each along the linear track 362.
  • one of the wedge faces 360 defines a recess 366.
  • one recess 366 may in use suffice to retain the locking feature 260 of the strip seal 200 and, thus, the strip seal 200 as a whole. That is to say, the recess 366 is configured to receive the locking member 260 of the strip seal 200 so that the locking member 260 may bend away from the main body 220, e.g. as shown in Figure 4 .
  • the recess 366 defines a section of the linear track 362 which an increased cross section.
  • the recess 366 may be provided so as to increase the depth of the linear track 360, measured along the circumferential direction.
  • the recess 366 may be provided so as to increase the width of the linear track 360, measured along the axial and/or radial direction.
  • the recess 366 is provided to expand the linear track 360 in the axial direction and the radial direction. Put another way, the recess 366 is an enlarged region of the track 360 for receiving the locking member 260.
  • Figure 8 is a partial perspective view of an annular assembly 400 of gas turbine. According to the example illustrated in Figure 6 , two annular segments 300 are arrayed to define the annular assembly 400, e.g. a partial stator vane.
  • the two annular segments 300 have been arranged next to one another so that a wedge face 360 of the first annular segment 300 and a corresponding wedge face 360 of the second annular segment 300 are interfaced.
  • a gap 410 is formed between the interfaced wedge faces 360 of the annular segments 300.
  • the strip seal 200 is configured to seal the gap 410 such that, in use, air is inhibited from passing through the gap.
  • the strip seal 200 is inserted into an opening 420 in an outer face of the (sub-)assembly 400.
  • the opening 420 is formed from the open ends 364 of the linear tracks 362 in the annular segments 300.
  • the linear tracks 362 co-operate to define a slot 430 extending into the assembly 400.
  • the locking member 260 is depressed by at least one of the annular segments 300. That is to say, insertion of the strip seal 200 causes a mechanical deformation of the strip seal 200, in that the locking member 260 is forced into the cavity 240 formed by the strip seal 200.
  • the locking member 260 As the strip seal 200 is inserted farther into the slot 420, the locking member 260 remains in the cavity 240 until the locking member 260 reaches the recess 226. The strip seal 260 then assumes its previous configuration, i.e. the locking member 260 extends from the main body 220 and into the recess 226.
  • the strip seal 200 may be recovered. Whilst the locking member 260 prevents the strip seal 200 from sliding out of the slot 430 through the opening 420, the strip seal 200 is removable from the annular segments 300 once these have been separated. Moreover, the strip seal 200 may be recoverable in a substantially undamaged condition and re-useable.
  • FIG. 8 is a stator vane assembly, although it will be understood sealing may be achieved in the same way between blade platforms of a rotor blade assembly.
EP18206450.1A 2018-11-15 2018-11-15 Streifendichtung, ringsegment und verfahren für eine gasturbine Withdrawn EP3653844A1 (de)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP18206450.1A EP3653844A1 (de) 2018-11-15 2018-11-15 Streifendichtung, ringsegment und verfahren für eine gasturbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP18206450.1A EP3653844A1 (de) 2018-11-15 2018-11-15 Streifendichtung, ringsegment und verfahren für eine gasturbine

Publications (1)

Publication Number Publication Date
EP3653844A1 true EP3653844A1 (de) 2020-05-20

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EP18206450.1A Withdrawn EP3653844A1 (de) 2018-11-15 2018-11-15 Streifendichtung, ringsegment und verfahren für eine gasturbine

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113503190A (zh) * 2021-09-13 2021-10-15 中国航发上海商用航空发动机制造有限责任公司 航空发动机静子叶片缘板封严条，静子叶片及封严结构

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1798380A2 (de) * 2005-12-16 2007-06-20 General Electric Company Turbinendüse mit Dichtstreifen
EP1832716A2 (de) * 2006-03-09 2007-09-12 United Technologies Corporation Segmentierte Komponentendichtung
EP2843197A2 (de) * 2013-08-29 2015-03-04 Alstom Technology Ltd Schaufel einer rotierenden Strömungsmaschine mit radialer Streifendichtung
US20150354381A1 (en) * 2013-02-05 2015-12-10 Snecma Flow distribution blading comprising an improved sealing plate
US20160032742A1 (en) * 2013-03-13 2016-02-04 United Technologies Corporation Stator segment

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1798380A2 (de) * 2005-12-16 2007-06-20 General Electric Company Turbinendüse mit Dichtstreifen
EP1832716A2 (de) * 2006-03-09 2007-09-12 United Technologies Corporation Segmentierte Komponentendichtung
US20150354381A1 (en) * 2013-02-05 2015-12-10 Snecma Flow distribution blading comprising an improved sealing plate
US20160032742A1 (en) * 2013-03-13 2016-02-04 United Technologies Corporation Stator segment
EP2843197A2 (de) * 2013-08-29 2015-03-04 Alstom Technology Ltd Schaufel einer rotierenden Strömungsmaschine mit radialer Streifendichtung

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113503190A (zh) * 2021-09-13 2021-10-15 中国航发上海商用航空发动机制造有限责任公司 航空发动机静子叶片缘板封严条，静子叶片及封严结构
CN113503190B (zh) * 2021-09-13 2022-02-15 中国航发上海商用航空发动机制造有限责任公司 航空发动机静子叶片缘板封严条，静子叶片及封严结构

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