WO2018100160A1 - Turbine à gaz - Google Patents

Turbine à gaz Download PDF

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Publication number
WO2018100160A1
WO2018100160A1 PCT/EP2017/081193 EP2017081193W WO2018100160A1 WO 2018100160 A1 WO2018100160 A1 WO 2018100160A1 EP 2017081193 W EP2017081193 W EP 2017081193W WO 2018100160 A1 WO2018100160 A1 WO 2018100160A1
Authority
WO
WIPO (PCT)
Prior art keywords
turbine blade
gas turbine
blade elements
disc
connecting means
Prior art date
Application number
PCT/EP2017/081193
Other languages
German (de)
English (en)
Inventor
Karl Schreiber
Original Assignee
Rolls-Royce Deutschland Ltd & Co Kg
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 Rolls-Royce Deutschland Ltd & Co Kg filed Critical Rolls-Royce Deutschland Ltd & Co Kg
Priority to EP17822557.9A priority Critical patent/EP3548704A1/fr
Priority to US16/462,466 priority patent/US20190376392A1/en
Publication of WO2018100160A1 publication Critical patent/WO2018100160A1/fr

Links

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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/284—Selection of ceramic materials
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3084—Fixing blades to rotors; Blade roots ; Blade spacers the blades being made of ceramics
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/32—Locking, e.g. by final locking blades or keys
    • F01D5/323—Locking of axial insertion type blades by means of a key or the like parallel to the axis of the 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • F05D2300/6033—Ceramic matrix composites [CMC]
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to a gas turbine having the features of claim 1.
  • Gas turbines such as aircraft engines or stationary gas turbines, are thermally and mechanically highly loaded units.
  • the efficiency of a gas turbine is greatly influenced by the thermal and mechanical strength of the gas turbine.
  • It is known in the art for example from WO 2012 160819 A1, to connect the turbine blades in a form-fitting manner via a so-called fir tree root with a turbine disk.
  • This positive connection requires a considerable volume of material and significantly affects the weight and the load capacity of the gas turbine.
  • compressor designs can be estimated that - compared to an integral blisk design - about 30% of the total weight of a stage for positive engagement are needed.
  • a blisk design is not sensible and not common due to very different material requirements of turbine blade and turbine disk. There is therefore the task of adapting gas turbines to the particular conditions of use.
  • the gas turbine has at least one disc, wherein turbine blade elements are connected via connecting means with the at least one disc.
  • the connecting means are arranged in the interior of the turbine blade elements, wherein the turbine blade elements are arranged in the radial extension in a region radially above the disc, in particular in a region are arranged, which is in the operation of the gas turbine in the driving air flow.
  • the turbine blade elements have at least two zones made of different materials, wherein the at least two zones adjoin one another in particular in the radial extension and that radially below the connecting means, a zone with a pressure stress suitable material, in particular a ceramic, in particular a yittrium-stabilized zirconium oxide, is arranged and radially above the connecting means, a zone with a zugêtsgeauchem material, in particular CMSX 4, is arranged.
  • the materials can be selected according to load.
  • at least one separating line can be arranged between the at least two zones of different material in the turbine blade elements radially below the connecting means.
  • the connecting means are thus in particular in a region of the turbine blade elements, which are exposed to the hot gas flow, i. in the aerodynamically effective area (aerofoil area) of the turbine blade elements.
  • the connecting means are arranged in a region in which comparatively small masses must be transmitted.
  • the cores for the turbine blade elements can be made integrally materially from the disc. Only in a radial area, which is in operation in the hot gas stream, is the positive connection to the enveloping Aerofoil Scheme. This leads to an approx. 70% reduction in mass, which must be transferred via the form fit and to a considerable reduction of stress in the aerofoil range. tensile stresses occur only in the blade region lying radially outside the positive connection. The area lying radially below is subjected to compressive stress.
  • gap control i.e., gap distance between blade tip and surrounding housing
  • thermal and elastic expansion due to design.
  • the connecting means are arranged radially inward, radially in the middle or radially outward in the region radially above the disk, in particular in the region of the turbine blade elements, which lies in the driving air flow during operation of the gas turbine.
  • the connecting means are formed positively, non-positively and / or cohesively.
  • Non-positive connection means may have a wedge connection, in particular in the disc and / or shrink joints to achieve a frictional connection.
  • Cohesive connection means may have a laser weld between the turbine blade elements and the disk.
  • wedge elements for bracing the disc with the turbine blade elements and / or for establishing a positive connection between the disc and the turbine blade elements.
  • the material of the core is then e.g. more elastic than the material of the wedge means.
  • the wedge means pushes the core e.g. against the inside of the turbine blade elements.
  • the turbine blade elements For connecting the turbine blade elements with the cores may be arranged radially outside a welded joint.
  • the turbine blade elements may consist of two parts, whereby only in the assembly of the parts the positive connections are made.
  • the gas turbines can be designed as an aircraft engine, as a vehicle drive, as a marine propulsion or stationary gas turbine.
  • the turbine blade elements or the disc can be particularly adapted and designed for use according to at least one of claims 1 to 15.
  • FIG. 1 shows a schematic representation of a gas turbine, here an aircraft turbine
  • FIG. 2A is a horizontal sectional view through a turbine blade element
  • FIG. 2B shows a sectional view through the turbine blade element according to FIG. 2A along the line A-A;
  • Fig. 3A shows an alternative embodiment of the turbine blade element with a
  • FIG. 3B the embodiment of FIG. 3A with a driven wedge means for
  • Fig. 4A shows an alternative embodiment of the turbine blade element with a
  • FIG. 4B shows an alternative embodiment of the turbine blade element with a
  • the individual components of the gas turbine 100 are arranged one behind the other along a rotation axis or center axis M, wherein the gas turbine 100 is designed as a turbofan engine.
  • the gas turbine 100 is designed as a turbofan engine.
  • air is sucked in along an inlet direction R by means of a fan F.
  • This arranged in a fan housing FC fan F is driven by a rotor shaft S, which is rotated by a turbine TT of the gas turbine 100 in rotation.
  • the turbine TT adjoins a compressor V, which has, for example, a low-pressure compressor 11 and a high-pressure compressor 12.
  • the fan F leads the compressor V and the Bypass channel B air to.
  • the bypass channel B in this case extends around a compressor V and the turbine TT comprehensive core engine, which includes a primary flow channel for the supplied through the fan F the core engine air.
  • the air conveyed into the primary flow passage via the compressor V enters a combustion chamber section BK of the core engine in which the driving power for driving the turbine TT is generated.
  • the turbine TT has a high-pressure turbine 13, a medium-pressure turbine 14 and a low-pressure turbine 15.
  • the low-pressure turbine 15 drives the rotor shaft S and thus the fan F via the energy released during combustion in order to generate the required thrust via the air conveyed into the bypass duct B.
  • Both the air from the bypass passage B and the exhaust gases from the primary flow passage of the core engine flow through an outlet A at the end of the engine T.
  • the outlet A in this case usually has a discharge nozzle with a centrally arranged outlet cone C.
  • Both in the area of the (axial) compressor with its low-pressure compressor 1 1 and its high-pressure compressor 12 and in the area of the turbine TT known rotating blade assemblies are used around the central axis M, each having a blade row and in which the blades on a ring or disc-shaped blade carrier are provided.
  • the ring-shaped or disk-shaped blade carrier can in principle be integrally bladed and thus be manufactured in bling or blisk construction. Alternatively, the fixing of individual blades via their respective blade root on a ring-shaped or disk-shaped blade carrier is possible.
  • FIG. 2A shows a horizontal sectional view through a blade, here a turbine blade element 1.
  • the turbine blade element 1 surrounds inside a core 4, which is integrally connected to a disc 5 of the turbine TT. This can be seen in more detail in the sectional view of FIG. 2B.
  • the disc 5 has the core 4, which projects into the interior of the turbine blade element 1.
  • the connection between the disk 5 and the turbine blade element 1 takes place here via a positive connection means 2 in the interior of the turbine blade element 1.
  • the connecting means 2 in the illustrated here Embodiment, a positive locking means 3, with which the turbine blade element 1 is fixed axially and / or radially.
  • a positive connection means 2 can be combined with non-positive and / or material connection means 2.
  • the positive locking means 3 is formed here on the radially outer edge of the core 4 as a mushroom-shaped formation which forms a shoulder. In the interior of the turbine blade element 1, a corresponding projection is formed which engages with the shoulder of the positive locking means 3 on the core 4.
  • the positive locking means 3 may for example also have an undercut.
  • a typical weight for a turbine blade element 1 in an aircraft engine is between 50 and 150 g. For stationary gas turbines, the weight can be significantly higher.
  • the tip of the turbine blade elements 1 extends radially away from the write 4 away over a height H.
  • the positive connection means 2 is located approximately at half the height of the blade height or on the half of the region H1 of the turbine blade element which is exposed to the hot, driving airflow L during operation of the gas turbine.
  • the turbine blade element 1 can be divided in radial extent into two zones Z1, Z2.
  • the first zone Z1 extends from the base of the turbine blade element 1 to the positive connection means 2.
  • the second zone Z2 extends from the positive connection means 2 to the blade tip. Between the zones Z1, Z2 runs in the illustrated embodiment, a dividing line T between different materials.
  • first zone Z1 below the form-locking connection means 2 in particular, compressive stresses act, so that particularly pressure-voltage-resistant materials can be used here.
  • An example of this is, for example, ceramics, in particular a yttrium-stabilized zirconium oxide or CMC (ceramic matrix composites).
  • CMC ceramic matrix composites
  • CMSX-4 monocrystalline material
  • Hf 0.07% by weight Hf.
  • This material is particularly temperature-resistant. In principle, however, other high temperature resistant superalloys in the second zone Z2 can be used.
  • the positive connection means 2 is arranged in the illustrated embodiment substantially at half the height of the turbine blade element 1.
  • the interlocking connection means 2 may also be closer to the base, i. closer to the disc 5 or closer to the tip of the turbine blade element 1, be arranged.
  • FIG. 2B also shows that cooling air from the region of the disk 5 can enter radially in the interior of the turbine blade element 1.
  • FIG. 3A shows a detail of one embodiment of a connection between a turbine blade element 1 and a disk 5.
  • the core 4 of the disk 5 in this case has a wedge means 7 that can be driven into a corresponding gap of the core 4.
  • Fig. 3A the not yet driven position of the wedge means 7 (the core 4 then has a small cross section) is shown. In this state, the turbine blade element 1 can be plugged.
  • a driven position of the wedge means 7 is shown, that is, the cross section of the core 4 increases, so that a positive connection means 2 between the core 4 and the turbine blade element 1 is produced.
  • FIGS. 4A and 4B A further alternative embodiment is shown in FIGS. 4A and 4B.
  • the wedge means 7 is used to produce a frictional connection means 2 between the core 4 of the disc 5.
  • Fig. 4A is a sectional view is shown, in which the core 4 is shown without the bracing wedge means 7.
  • the core 4 has only a prefabricated gap (dashed lines shown here), in which the tel 7 can be used. In the position shown here, there is no connection on the side walls between the turbine blade elements 1 and the core 4.
  • Such a connection can be combined, for example, with a positive locking means 3 (as in the embodiments shown in FIGS. 3A, 3B). It is also possible to produce an integral connection alternatively or additionally.
  • the configuration according to FIGS. 3 and 4 can be produced such that sleeve-shaped turbine blade elements 1 are placed over the core 4, ie the turbine blade elements 1 are open at the radially outer end. After fitting, then the positive connection, the frictional connection and / the material connection can be made.
  • the turbine blade elements 1 are each formed closed on the radially outer edge. This can be done, for example, by welding a cover after the connection has been made, as described above. However, it is also possible for the radially outer end to remain open, so that cooling air K entering below into the turbine blade elements 1 can exit at the top. LIST OF REFERENCE NUMBERS

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Ceramic Engineering (AREA)
  • Architecture (AREA)
  • Materials Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une turbine à gaz comprenant au moins un disque (5), des éléments d'aube de turbine (1) étant reliés audit au moins un disque (5) par le biais de moyens de liaison (2), caractérisée en ce que les moyens de liaison (2) sont disposés dans une région (H) radialement au-dessus du disque (5) dans la direction radiale des éléments d'aube de turbine (1) à l'intérieur des éléments d'aube de turbine (1), en particulier dans une région (H1) qui se situe dans le flux d'air (L) d'entraînement lors du fonctionnement de la turbine à gaz (100) et les éléments d'aube de turbine (1) comprennent au moins deux zones (Z1, Z2) constituées de matériaux différents, lesdites au moins deux zones (Z1, Z2) se raccordant les unes aux autres en particulier dans la direction radiale, et en ce qu'une zone (Z1) comprenant un matériau approprié pour les contraintes de compression, en particulier une céramique, en particulier de l'oxyde de zirconium stabilisé à l'yttrium, est disposée radialement en dessous des moyens de liaison (2) et une zone (Z2) comprenant un matériau approprié pour les contraintes de traction, en particulier CMSX-4, est disposée radialement au-dessus des moyens de liaison (2).
PCT/EP2017/081193 2016-12-01 2017-12-01 Turbine à gaz WO2018100160A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP17822557.9A EP3548704A1 (fr) 2016-12-01 2017-12-01 Turbine à gaz
US16/462,466 US20190376392A1 (en) 2016-12-01 2017-12-01 Gas turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102016123248.3A DE102016123248A1 (de) 2016-12-01 2016-12-01 Gasturbine
DE102016123248.3 2016-12-01

Publications (1)

Publication Number Publication Date
WO2018100160A1 true WO2018100160A1 (fr) 2018-06-07

Family

ID=60888356

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2017/081193 WO2018100160A1 (fr) 2016-12-01 2017-12-01 Turbine à gaz

Country Status (4)

Country Link
US (1) US20190376392A1 (fr)
EP (1) EP3548704A1 (fr)
DE (1) DE102016123248A1 (fr)
WO (1) WO2018100160A1 (fr)

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1517004A1 (fr) * 2003-09-19 2005-03-23 Snecma Moteurs Roue de turbine pour turbomachine et procédé de montage d'une telle roue
EP2469031A2 (fr) * 2010-12-27 2012-06-27 General Electric Company Composants de surface portante de turbine contenant des matériaux à base de céramique et procédés associés
WO2012160819A1 (fr) 2011-05-23 2012-11-29 株式会社 東芝 Aube de rotor de turbine et turbine à vapeur
FR2995933A1 (fr) * 2012-09-26 2014-03-28 Snecma Aube pour turbomachine en materiau composite a pied en forme de bulbe

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2553078A (en) * 1947-03-29 1951-05-15 United Aircraft Corp Turbine blade mounting
DE3539903A1 (de) * 1985-11-11 1987-05-14 Kloeckner Humboldt Deutz Ag Gasturbine mit einem keramischen laufrad
EP1329592A1 (fr) * 2002-01-18 2003-07-23 Siemens Aktiengesellschaft Turbine avec au moins quatre stages et utilisation des aubes en masse réduite

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1517004A1 (fr) * 2003-09-19 2005-03-23 Snecma Moteurs Roue de turbine pour turbomachine et procédé de montage d'une telle roue
EP2469031A2 (fr) * 2010-12-27 2012-06-27 General Electric Company Composants de surface portante de turbine contenant des matériaux à base de céramique et procédés associés
WO2012160819A1 (fr) 2011-05-23 2012-11-29 株式会社 東芝 Aube de rotor de turbine et turbine à vapeur
FR2995933A1 (fr) * 2012-09-26 2014-03-28 Snecma Aube pour turbomachine en materiau composite a pied en forme de bulbe

Also Published As

Publication number Publication date
EP3548704A1 (fr) 2019-10-09
DE102016123248A1 (de) 2018-06-07
US20190376392A1 (en) 2019-12-12

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