EP2824279A1 - Turbomachine, aube directrice et aube mobile - Google Patents

Turbomachine, aube directrice et aube mobile Download PDF

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Publication number
EP2824279A1
EP2824279A1 EP13175726.2A EP13175726A EP2824279A1 EP 2824279 A1 EP2824279 A1 EP 2824279A1 EP 13175726 A EP13175726 A EP 13175726A EP 2824279 A1 EP2824279 A1 EP 2824279A1
Authority
EP
European Patent Office
Prior art keywords
ring
blade
axial
row
overhang
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.)
Granted
Application number
EP13175726.2A
Other languages
German (de)
English (en)
Other versions
EP2824279B1 (fr
Inventor
Christoph Lauer
Markus Hirschmann
Yannick Dr. Muller
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines 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 MTU Aero Engines AG filed Critical MTU Aero Engines AG
Priority to EP13175726.2A priority Critical patent/EP2824279B1/fr
Priority to ES13175726T priority patent/ES2724533T3/es
Publication of EP2824279A1 publication Critical patent/EP2824279A1/fr
Application granted granted Critical
Publication of EP2824279B1 publication Critical patent/EP2824279B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D11/001—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
    • 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
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/127—Vortex generators, turbulators, or the like, for mixing

Definitions

  • the invention relates to a turbomachine according to the preamble of patent claim 1, a vane and a blade.
  • a turbomachine is shown with a sealing device which is formed as a labyrinth seal between a radially outer stator ring and a radially inner rotor ring.
  • the rings are in axial overlap, wherein the rotor ring in addition to its radially outwardly directed end portion to approximately its half axial extent has a circumferential radially outwardly directed sealing ridge.
  • the stator ring has a radially inwardly directed end portion.
  • FIG. 2 shows a turbomachine with a sealing device between a stationary airfoil row arranged in a turbine intermediate housing and a subsequent blade row.
  • the object of the invention is to provide a turbomachine having an alternative sealing means for sealing an axial gap between a series of airfoils and a downstream blade row.
  • the sealing device has a downstream-oriented guide vane-side axial ring, which extends axially beyond a platform ring trailing edge of the airfoil row and has a radially outwardly directed end portion.
  • the sealing device has an upstream platform ring overhang, which is arranged on the blade row and having a radially inwardly directed end portion, wherein the axial ring is arranged radially inward to the platform ring overhang and forms an axial overlap with this.
  • the front platform ring overhang follows, at least in sections, an ideal flow pattern of the hot gas over the rotor blade-side platform ring.
  • the formation of the airfoil series without a downstream or rear platform ring overhang in conjunction with the radially inwardly running rear hook-like axial ring and the axially outer overlap with the Axialring located radially outer front hook-like platform ring overhang causes a sealing space or a vortex zone in which / the incoming hot gas with a high tangential velocity component is applied.
  • the inflowing hot gas is swirled and thereby made much more difficult or prevented at its entry into the cooling channel.
  • the sealing space acts as a damping area between the hot gas duct and the inner cooling channel, whereby pressure differences between the hot gas duct and the inner cooling channel are weakened, whereby also a hot gas from the hot gas duct into the inner cooling channel is difficult.
  • the sealing device according to the invention allows large axial and radial relative movements of the rotor and the stator to each other in comparison to conventional Fischmauldichtitch at a small axial gap. Due to the fact that the front platform ring overhang follows, at least in sections, an ideal flow course of the hot gas over the blade row side platform ring, in the region of the platform overhang in the hot gas channel, eddies are prevented or substantially reduced.
  • the sealing device is preferably arranged on the turbine side of the turbomachine such as a gas turbine and in particular an aircraft engine.
  • the airfoil series is a vane row in a low-pressure turbine of the turbomachine.
  • the airfoil series can also consist of a plurality of blade profiles arranged in a turbine intermediate housing between a high-pressure turbine and a low-pressure turbine.
  • the end portions are spaced apart in the radial direction. As a result, a radial gap is essentially created, can be blown through the cooling air from the cooling passage in the sealing chamber.
  • the end portions terminate at the same radial height.
  • an axial gap is substantially created, can be blown through the cooling air from the cooling passage in the sealing chamber.
  • the end portion of the axial ring has a greater radial extent than the end portion of the front platform ring overhang.
  • the end portion of the rotating platform ring overhang is shortened in this embodiment compared to the end portion of the fixed axial ring.
  • the end portions of the platform ring overhang are stabilized by the short radial extent, which prevents the introduction of disturbing vibrations in the rotor by the rotating end portions.
  • the end portions are parallel and oriented radially offset from one another.
  • the end portions are orthogonal to the axial direction of the turbomachine. Due to the parallelism, an axial component of the radial gap is kept constant despite different thermal radial expansion behavior of the axial ring and the platform ring overhang or of the stator and of the rotor.
  • the end portions are set to each other.
  • the end portion of the axial ring is inclined downstream in the axial direction and the end portion of the front platform ring overhang is orthogonal to the axial direction. Due to the inclination or inclination of the gap between the cooling chamber and the sealing space can be opened.
  • a guide vane according to the invention for a turbomachine according to the invention has a downstream-oriented axial overhang which extends axially beyond a platform trailing edge and has a radially outwardly directed end portion.
  • Such a guide vane allows the formation of a row of guide vanes into which the preferred sealing device for preventing fuel gas is integrated in sections.
  • a blade according to the invention for a turbomachine according to the invention has an upstream front platform overhang, which has a radially inwardly directed end portion. Such a blade allows the formation of a blade row, in which the preferred sealing means for preventing Schugaseizugs is partially integrated.
  • FIG. 1 a longitudinal section through a turbomachine in the region of a flow profile row 1 and a row of airfoils 1 downstream flow profile row 2 is shown.
  • the flow profile row is a row of guide vanes.
  • the guide blade row 1 and the blade row 2 have a plurality of blades 35, 36, which are arranged side by side in the circumferential direction of the turbomachine and their respective blade 4, 6 is arranged in the hot gas duct 8 of the turbomachine.
  • the turbomachine is a gas turbine and in particular an aircraft engine.
  • the guide vane row 1 and the blade row 2 are arranged in particular on the turbine side and, for example, in the low-pressure turbine of the turbomachine.
  • the hot gas channel 8 is from a hot gas stream as shown in FIG. 2 flows through from left to right. It is bounded radially on the inside by a guide-row-side platform ring 10 and by a rotor blade-side platform ring 12. Radially inside the hot gas duct 8, a cooling space 14 is formed, which is flowed through by a cooling air flow.
  • a sealing device 18 is arranged in the axial gap between the guide vane row 1 and the blade row 2.
  • FIG. 1 shown first embodiment of the sealing device 18 forms a ring-like sealing space 20.
  • the sealing space 20 is in the in FIG. 1 shown embodiment via an annular axial outer gap 21 to the hot gas channel 8 and via an annular axial inner gap 23 to the cooling channel 14.
  • the hot gas partial stream 16 enters the sealing space 20 and is swirled in this.
  • a partial cooling air flow 22nd injected, which lays on a hot gas partial flow vortex 25 in the sealing chamber 20 and exits into the hot gas channel 8.
  • the sealing device 18 has an opposite to a guide vane row-side front platform ring trailing edge 24 radially inwardly disposed axial ring 26 which is oriented downstream in the direction of the blade row 2 and has a radially outwardly directed end portion 28.
  • the sealing device 18 has a blade row-side front platform ring overhang 30, which is oriented downstream in the direction of the guide blade row 1 and has a radially inwardly directed end portion 32.
  • the circumferential end portions 28, 32 thus have in opposite radial directions, wherein they face each other.
  • the axial ring 26 and the platform ring overhang 30 are in axial overlap. They are in the in FIG. 1 shown embodiment in axial overlap that the sealing space 20 is not open over its entire axial extent to the hot gas channel 8, but only in the area near the platform ring trailing edge 24th
  • the axial ring 26 extends with a cylindrical annular portion 34 in the axial direction of the turbomachine and is arranged radially inward to the platform ring overhang 30. Its end portion 28 extends from the ring portion 34 and is made inclined downstream, extending radially outwardly. The end portion 28 thus points in the direction of the front platform ring overhang 30.
  • the front platform ring overhang 30 follows in the embodiment shown with a ring portion 37 an ideal flow of the hot gas over the blade row side platform ring 12. He goes for this purpose on a stepless in the platform 12 and has viewed in the flow direction a funnel-shaped opening contour.
  • the end portion 32 of the front platform ring overhang 30 extends from the ring portion 37. It is orthogonal to the axial direction and thus orthogonal to the ring portion 34 of the axial ring 26, extending radially inwardly.
  • the end portion 32 has a short Radial extent than the end portion 28 of the axial ring 26. In the illustrated embodiment, it has such a radial extent that it ends at or approximately at the same radial height as the end portion 28 of the axial ring 26. Downstream forms the end portion 32 of the platform ring overhang 30 together with the end portion 28 of the axial ring 26, the inner gap 23, the here in FIG. 1 almost an axial gap.
  • FIG. 2 shown embodiment of a sealing device 18 between a stator blade 1 and a downstream blade row 2 of a turbomachine to prevent a hot gas inlet 16 from a radially outer hot gas channel 8 in a radially inner cooling channel 14 to form a sealing space 20 are essentially different from the embodiment according to FIG. 1 both end portions 28, 32 oriented orthogonal to the axial ring portion 34 and to the axial direction of the turbomachine.
  • the end sections 28, 32 are aligned almost parallel to one another.
  • a ring-like inner gap 23 for injecting a cooling air partial flow 22 into the sealing space 20 has a greater axial component than a radial component and thus is virtually a radial gap.
  • end sections 28, 32 are substantially different from a guide blade-side platform ring trailing edge 24 in that a ring-like axial outer gap 21 extends virtually over the entire axial extent of the axial ring 26 and the sealing space 20 thus opens to the hot gas channel 8 virtually over its entire axial extent is.
  • a respective vane 35 of the vane row 1 is formed without a rear platform overhang.
  • the vane 35 has in each case one to a rear platform edge radially inwardly positioned axial projection with a radially outwardly directed end portion.
  • the adjacent axial projections form the axial ring 26, wherein the end regions form the peripheral end section 28.
  • FIG. 3 a blade 36 according to the invention for the formation of an aforementioned blade row 2 is shown.
  • the blade 36 has seen in the direction of flow of a hot gas, a front platform overhang 38 with a radially inwardly directed end portion 40.
  • the front platform overhang 38 follows an ideal flow of the hot gas through a paddle platform 42.
  • the front platform overhang 38 is steplessly into the paddle platform 42 and is in accordance with the blade platform 42 ascending inclined in the flow direction.
  • the in FIG. 3 shown blade 36 a rear platform overhang 44 which is aligned almost axially to the axial direction.
  • the adjacent front platform overhangs 38 form the front platform ring overhang 30, wherein the end portions 40 form the peripheral end portion 32.
  • a turbomachine having sealing means for preventing hot gas ingress through an axial gap between a series of airfoils and a blade row from a hot gas passage into a cooling passage, the sealing means having a downstream orifice-side axial ring extending axially beyond a platform ring trailing edge of the airfoil row and a radial one has an outwardly directed end portion, and an upstream blade-side platform ring overhang having a radially inwardly directed end portion, wherein the Axialring is arranged radially inwardly to the platform ring overhang and forms an axial overlap therewith, wherein the front platform ring overhang at least partially an ideal flow of the Hot gas over the blade row side platform ring follows, as well as a vane and a blade.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13175726.2A 2013-07-09 2013-07-09 Turbomachine avec dispositif d'étanchéité Active EP2824279B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP13175726.2A EP2824279B1 (fr) 2013-07-09 2013-07-09 Turbomachine avec dispositif d'étanchéité
ES13175726T ES2724533T3 (es) 2013-07-09 2013-07-09 Turbomáquina con estructura de sellado

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP13175726.2A EP2824279B1 (fr) 2013-07-09 2013-07-09 Turbomachine avec dispositif d'étanchéité

Publications (2)

Publication Number Publication Date
EP2824279A1 true EP2824279A1 (fr) 2015-01-14
EP2824279B1 EP2824279B1 (fr) 2019-04-03

Family

ID=48748003

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13175726.2A Active EP2824279B1 (fr) 2013-07-09 2013-07-09 Turbomachine avec dispositif d'étanchéité

Country Status (2)

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EP (1) EP2824279B1 (fr)
ES (1) ES2724533T3 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111287803A (zh) * 2018-12-07 2020-06-16 安萨尔多能源公司 用于燃气涡轮的定子组件和包括所述定子组件的燃气涡轮

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2847012A1 (de) 1977-10-31 1979-05-03 Gen Electric Kombinierte turbinenmantel- und schaufelhalterung
EP1380726A2 (fr) * 2002-07-10 2004-01-14 Mitsubishi Heavy Industries, Ltd. Aube statorique pour turbine à gaz et turbine à gaz comprenant cet élément
US7540709B1 (en) 2005-10-20 2009-06-02 Florida Turbine Technologies, Inc. Box rim cavity for a gas turbine engine
US20100172749A1 (en) 2007-03-29 2010-07-08 Mitsuhashi Katsunori Wall of turbo machine and turbo machine
US20100183426A1 (en) * 2009-01-19 2010-07-22 George Liang Fluidic rim seal system for turbine engines
US8075256B2 (en) 2008-09-25 2011-12-13 Siemens Energy, Inc. Ingestion resistant seal assembly
EP2573329A2 (fr) 2011-09-22 2013-03-27 Pratt & Whitney Canada Corp. Architecture de système d'air pour un module de cadre de turbine intermédiaire

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2847012A1 (de) 1977-10-31 1979-05-03 Gen Electric Kombinierte turbinenmantel- und schaufelhalterung
EP1380726A2 (fr) * 2002-07-10 2004-01-14 Mitsubishi Heavy Industries, Ltd. Aube statorique pour turbine à gaz et turbine à gaz comprenant cet élément
US7540709B1 (en) 2005-10-20 2009-06-02 Florida Turbine Technologies, Inc. Box rim cavity for a gas turbine engine
US20100172749A1 (en) 2007-03-29 2010-07-08 Mitsuhashi Katsunori Wall of turbo machine and turbo machine
US8075256B2 (en) 2008-09-25 2011-12-13 Siemens Energy, Inc. Ingestion resistant seal assembly
US20100183426A1 (en) * 2009-01-19 2010-07-22 George Liang Fluidic rim seal system for turbine engines
EP2573329A2 (fr) 2011-09-22 2013-03-27 Pratt & Whitney Canada Corp. Architecture de système d'air pour un module de cadre de turbine intermédiaire

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111287803A (zh) * 2018-12-07 2020-06-16 安萨尔多能源公司 用于燃气涡轮的定子组件和包括所述定子组件的燃气涡轮
CN111287803B (zh) * 2018-12-07 2023-07-14 安萨尔多能源公司 用于燃气涡轮的定子组件和包括所述定子组件的燃气涡轮

Also Published As

Publication number Publication date
EP2824279B1 (fr) 2019-04-03
ES2724533T3 (es) 2019-09-11

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