EP2159382A1 - Aube directrice pour turbine à gaz - Google Patents

Aube directrice pour turbine à gaz Download PDF

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Publication number
EP2159382A1
EP2159382A1 EP08015145A EP08015145A EP2159382A1 EP 2159382 A1 EP2159382 A1 EP 2159382A1 EP 08015145 A EP08015145 A EP 08015145A EP 08015145 A EP08015145 A EP 08015145A EP 2159382 A1 EP2159382 A1 EP 2159382A1
Authority
EP
European Patent Office
Prior art keywords
vane carrier
guide vane
rings
carrier
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Withdrawn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
EP08015145A
Other languages
German (de)
English (en)
Inventor
Francois Dr. Benkler
Andreas Dr. Böttcher
Tino Etzold
Daniel Grundei
Uwe Lohse
Ekkehard Dr. Maldfeld
Oliver Dr. Schneider
Shilun Dr. Sheng
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 EP08015145A priority Critical patent/EP2159382A1/fr
Priority to PCT/EP2009/060755 priority patent/WO2010023150A1/fr
Publication of EP2159382A1 publication Critical patent/EP2159382A1/fr
Withdrawn legal-status Critical Current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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/14—Casings modified therefor
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • F05D2240/14—Casings or housings protecting or supporting assemblies within

Definitions

  • the invention relates to a guide vane carrier for a gas turbine, comprising a tubular wall having an inlet end and an upstream end opposite the inflow end for an interior of the vane carrier in a flow path of the gas turbine fluid flowing, wherein in the wall at least one cooling channel for a coolant is provided which extends from a coolant inlet to a coolant outlet, respectively.
  • Turbinenleitschaufela of the type mentioned for a stationary gas turbine nowadays have in principle two identical half-tube-like support elements, which flange-together with each other screwed form a tubular shape. On the inner surfaces along the circumference extending grooves for receiving and mounting of vanes of the turbine unit of the gas turbine are provided.
  • the turbine unit of the gas turbine also has a flow path which increases in cross-section along the axial extent, whereby its outside diameter also increases continuously.
  • the turbine guide vane carrier of such a gas turbine is also designed to correspond: with a diameter which increases continuously in the flow direction on average. He is thus conical.
  • a guide blade carrier of the type mentioned at the beginning always has an inflow-side end and an outflow-side end for the hot gas.
  • the inflow-side end of the turbine guide vane carrier which is annular in cross-section has a smaller radius than the radius of the outflow end.
  • the hot gas generated in the combustion chamber is guided into the annular flow path of the gas turbine, which extends along the axial direction through the interior of the turbine vane support.
  • the flow path guides the injected hot gas to the downstream end of the turbine vane carrier. Meanwhile, the hot gas in the flow path relaxes working on the turbine blades and gives off heat to its environment. In addition, the temperature of the hot gas is further reduced by the supply of cooling air. As is known, this results in higher inlet temperatures than outlet temperatures for the hot gas.
  • the object of the invention is therefore to provide a guide vane carrier for a gas turbine, which may continue to be made of a comparatively inexpensive material at further elevated temperature conditions.
  • the invention provides that the guide blade carrier - viewed in the radial direction - is formed multi-layered. Due to the multiple layers of the guide blade carrier, each layer can be manufactured from a material that is adapted to its subarea, which is matched in each case to the mechanical and / or thermal stresses actually occurring. Due to this modular construction of the guide vane carrier, less highly alloyed and thus less costly cast steel grades can be used in the less hot regions of the vane carrier. This results in a reduction in material costs due to the use of a relatively inexpensive material for a particular area of the vane carrier.
  • the layers are formed from at least two concentrically arranged rings.
  • the rings lie against one another in a contact plane, wherein at least one of the rings has at least one groove arranged in its contact surface for forming the cooling channel.
  • each of the rings is manufactured separately.
  • the coolant can thus be guided in a simple manner along the axial extent of the guide blade carrier to the regions subjected to particularly high stress, without the production of the cooling channels requiring particular effort. Compared to an exclusively cast version of a guide vane carrier complex and expensive core geometries are avoided for any cooling channels.
  • At least one of the rings is formed by two half-rings.
  • the division of the rings in two half-rings is particularly advantageous for stationary gas turbines whose structural design - apart from the rotor - is divided by a dividing plane into a lower half and an upper half.
  • the rings according to the aforementioned embodiment are adapted to the structural design of the stationary gas turbine.
  • the guide blade carrier is formed from two semi-tubular guide blade carrier elements, which are screwed together in a flange manner.
  • the ring is not only divided into two half rings. Rather, it can also be divided into more than two segments, which significantly simplifies the production of half-rings. To produce the half rings, the segments can then be welded together.
  • the outermost layer of the guide blade carrier is unsegmented, apart from the division into two half rings.
  • the half-ring consisting of the segmented plates can be encompassed by an outer ring produced in one piece by casting, which consists of a low-alloy cast steel.
  • More internally located layers of the vane carrier can then consist of heat-resistant materials.
  • the thermal expansion of the turbine guide vane carrier can be positively influenced by the suitable choice of the different steel alloys.
  • the outermost layer thus serves as a shell for the radially inner layers of the turbine guide vane carrier.
  • two adjacent in the semi-circular contact plane half-rings are connected to each other by hot isostatic pressing.
  • hot isostatic pressing segments of different steel alloys can be connected to each other in a particularly simple manner.
  • Alternative joining methods are also applicable.
  • a guide blade carrier composed of a plurality of layers can have a multiplicity of cooling channels distributed along its circumference in order to keep the temperatures occurring in its material below a maximum permissible operating temperature. Due to the distribution of the cooling channels along the circumference of the guide blade carrier can be cooled during operation particularly evenly with cooling air, which avoids different thermal loads and different thermal expansions. Thus, a particularly efficient cooling of the guide blade carrier can also take place, in particular in regions near the flange, so that the ovalization effects that otherwise occur there are at least reduced, if not even avoided.
  • circumferential grooves for fastening guide vanes of a turbine unit or a compressor unit arranged in a ring are arranged on the inner surface of the guide blade carrier.
  • a guide vane carrier designed as a turbine guide vane carrier is preferably located in a pressure jacket of the gas turbine designed as an outer casing, in which a fuel is burnt to a hot gas for driving the turbine. If, on the other hand, the vane carrier is a compressor vane carrier is formed, this is flowed through instead of the hot gas of compressed air. In this case, the guide vane carrier is also cylindrical rather than conical.
  • a vane support according to the invention represents a composite construction in a multi-layer component construction, each of which is composed of welded and / or cast plate-shaped subcomponents. Depending on the local position and the operating temperatures occurring in this local position, different materials for the individual layers can be selected. As a result, in the region of high operating temperatures, the plate segments connected by hot isostatic pressing can be made of heat-resistant, weldable steel. The plate segments are joined together by welding in half rings, which form the inner part of the guide vane carrier. Subsequently, consisting of the multi-layer plates half rings are each inserted in a half outer ring, which is formed from a low alloy steel casting unsegmented in one piece. A guide vane carrier according to the invention can thus withstand increased limit application temperatures permanently and reliably.
  • FIG. 1 shows the cross section through a guide vane 10 of a stationary, axially flow-through gas turbine.
  • the vane carrier 10 can be arranged both in a turbine of the gas turbine and / or in the compressor of the gas turbine and accordingly be designed as a turbine guide vane carrier and / or as a compressor vane carrier.
  • the turbine vane carrier 10 is formed entirely of two half-tube-like vane support members 12, of which only the upper in FIG. 1 is shown.
  • the lower, not shown Leitschaufelaelement is identical to the upper Leitschaufelaelement 12, wherein both Leitschaufelanymaschine 12 in a parting plane 15 abut against each other and by means of the flanges 17 extending screws are firmly connected.
  • the turbine vane carrier 10 has an axial extent which is formed perpendicular to the plane of the drawing.
  • the turbine guide vane carrier 10 may be conical, for example, a compressor vane carrier is rather cylindrical.
  • the tubular turbine vane carrier 10 has a first end and a second end opposite the first end.
  • the inner diameter of the first end is substantially smaller than the inner diameter of the second end.
  • the turbine guide carrier 10 is arranged in a pressure jacket of the gas turbine such that the first end of the outlet a combustion chamber not shown opposite.
  • the second end of the turbine vane carrier 10 is located opposite an output diffuser, not shown, of the gas turbine.
  • annular flow path In an inner space 14 enclosed by the turbine vane support 10, there extends in the axial direction an annular flow path, which is bounded radially outward and radially inward by suitable elements, ie platforms of turbine blades and guide rings. For the sake of clarity, these elements are in FIG. 1 not shown, as well as within the turbine nozzle carrier 10 in cross-sectionally configured annular flow path for the hot gas generated in the combustion chamber.
  • the first end of the turbine vane support is the inflow-side end and the second end is the outflow-side end.
  • the turbine vane carrier 10 is principally formed by the semicircular walls 16 of the vane carrier elements 12. According to the invention, the wall 16 - viewed in the radial direction - formed in multiple layers. In the assembled turbine vane carrier 10, the layers are formed of at least two concentrically arranged rings 18a, 18b, 18c, which, as shown in FIG FIG. 1 only half are shown.
  • the wall 16 includes the outer first ring 18a, radially inwardly adjacent the second ring 18b, and radially inwardly disposed the third ring 18c. Due to the half-construction of an axially flow-through stationary gas turbine, the rings 18a, 18b, 18c are formed as half-rings, which have been connected by a hot isostatic pressing to the vane support member 12.
  • the lateral surfaces of the rings 18 or half rings are in each case in a circular arc-shaped contact plane 20 to each other, wherein at least one of the rings 18 or half rings has at least one arranged in its contact surface groove 22 for forming one or more cooling channels.
  • the two inner rings 18b, 18c are - viewed in the circumferential direction - formed of a plurality of segments 24, which are welded together at their joints.
  • each ring 18b, 18c and half ring 18 is a curved Plate which is composed of several segments 24.
  • the outer unsegmented half ring 18a engages around the two curved half rings 18b, 18c, the former being integrally formed from a low alloy steel casting.
  • the outer ring 18a is thus unsegmented and thus serves as a pressure resistant sheath around the turbine vane carrier 10.
  • the rings 18b and 18c as well as the outer ring 18a may be made of different steel alloys adapted to the respective thermal stress requirements. Since higher temperatures occur inside the turbine vane carrier 10 than in the outer region of the turbine vane carrier 10, it is advantageous to manufacture the inner rings 18c, 18d from a more heat-resistant material than the outer ring 18a. Thus, in the less hot regions of the turbine vane carrier 10, less expensive and less high alloyed steel castings may be used than in the hotter regions.
  • the cooling channels 22 arranged in the contact planes 20 can be introduced in a particularly simple manner by simply milling into the lateral surface of the segments 24 or the half rings 18b, 18c. After the introduction of the cooling channels, the mutually adjacent half rings are joined together by hot isostatic pressing to form a two-layer half ring, which then form an inner part of the turbine guide vane carrier 10.
  • the arrangement of the cooling channels and the supply of coolant through inlets not shown, the inner rings 18b, 18c can be cooled very easily and efficiently, so that they can permanently withstand the temperatures occurring during operation of the gas turbine.
  • segmented half rings 18b, 18c is also conceivable that they are also unsegmented, ie formed in one piece. This embodiment of the unsegmented half rings 18b, 18c is in FIG. 2 shown.
  • the two half rings 18b, 18c are also in the contact plane 20 to each other and are permanently connected together by the hot isostatic pressing.
  • cooling channels 22 arranged in the half rings or segments in a straight line along the axial direction A of the turbine guide vane carrier 10. Since the cooling channels 22 can be introduced into the lateral surface, ie contact surface, by simple mechanical stressing methods, any complex structures or meandering forms of cooling channels 22 are also conceivable, as they are described in US Pat FIG. 4 are shown as examples.
  • the invention relates to a guide rail carrier 10 for an axially flowed stationary gas turbine, comprising a tubular wall 16 with an inflow-side end and an inflow-side end opposite inflow end for flowing in the interior of the vane carrier in a flow path of the gas turbine fluid, wherein in the wall 16 at least a cooling channel 22 is provided for a coolant.
  • a guide blade carrier 10 which is suitable for particularly high operating temperatures and still relatively inexpensive to produce, it is proposed that the guide blade carrier 10 - viewed in the radial direction - is formed multi-layered.
  • the different layers of the vane carrier 10 may be interconnected by hot isostatic pressing, wherein the inner layers of the vane carrier 10 may be made of a high temperature resistant material, whereas outer layers of the vane carrier 10 may be made of a less temperature resistant material. Due to the multi-layer design of the guide vane carrier 10 also particularly simple cooling channels 22 inside the wall 16 of the vane carrier 10 can be produced.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP08015145A 2008-08-27 2008-08-27 Aube directrice pour turbine à gaz Withdrawn EP2159382A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP08015145A EP2159382A1 (fr) 2008-08-27 2008-08-27 Aube directrice pour turbine à gaz
PCT/EP2009/060755 WO2010023150A1 (fr) 2008-08-27 2009-08-20 Support d'aubes directrices pour une turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP08015145A EP2159382A1 (fr) 2008-08-27 2008-08-27 Aube directrice pour turbine à gaz

Publications (1)

Publication Number Publication Date
EP2159382A1 true EP2159382A1 (fr) 2010-03-03

Family

ID=40251810

Family Applications (1)

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EP08015145A Withdrawn EP2159382A1 (fr) 2008-08-27 2008-08-27 Aube directrice pour turbine à gaz

Country Status (2)

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EP (1) EP2159382A1 (fr)
WO (1) WO2010023150A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9062561B2 (en) 2010-09-29 2015-06-23 Rolls-Royce Plc Endwall component for a turbine stage of a gas turbine engine

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2930307A1 (fr) 2014-04-09 2015-10-14 Alstom Technology Ltd Support d'ailette de redresseur pour un compresseur ou une section de turbine d'une turbomachine axiale

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4431373A (en) * 1980-05-16 1984-02-14 United Technologies Corporation Flow directing assembly for a gas turbine engine
GB2261708A (en) * 1991-11-20 1993-05-26 Snecma Turbo-shaft engine casing and blade mounting
GB2412949A (en) * 2004-04-05 2005-10-12 Snecma Moteurs Turbine stator casing formed by hot isostatic compression
US7013652B2 (en) * 2001-11-20 2006-03-21 Alstom Technology Ltd Gas turbo set

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4431373A (en) * 1980-05-16 1984-02-14 United Technologies Corporation Flow directing assembly for a gas turbine engine
GB2261708A (en) * 1991-11-20 1993-05-26 Snecma Turbo-shaft engine casing and blade mounting
US7013652B2 (en) * 2001-11-20 2006-03-21 Alstom Technology Ltd Gas turbo set
GB2412949A (en) * 2004-04-05 2005-10-12 Snecma Moteurs Turbine stator casing formed by hot isostatic compression

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9062561B2 (en) 2010-09-29 2015-06-23 Rolls-Royce Plc Endwall component for a turbine stage of a gas turbine engine

Also Published As

Publication number Publication date
WO2010023150A1 (fr) 2010-03-04

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