EP0313826B1 - Turbine à gaz avec flux axial - Google Patents

Turbine à gaz avec flux axial Download PDF

Info

Publication number
EP0313826B1
EP0313826B1 EP88115694A EP88115694A EP0313826B1 EP 0313826 B1 EP0313826 B1 EP 0313826B1 EP 88115694 A EP88115694 A EP 88115694A EP 88115694 A EP88115694 A EP 88115694A EP 0313826 B1 EP0313826 B1 EP 0313826B1
Authority
EP
European Patent Office
Prior art keywords
blade
cooling
rotor
ring
air
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.)
Expired - Lifetime
Application number
EP88115694A
Other languages
German (de)
English (en)
Other versions
EP0313826A1 (fr
Inventor
Franz Kreitmeier
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.)
BBC Brown Boveri AG Switzerland
Original Assignee
BBC Brown Boveri AG Switzerland
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 BBC Brown Boveri AG Switzerland filed Critical BBC Brown Boveri AG Switzerland
Publication of EP0313826A1 publication Critical patent/EP0313826A1/fr
Application granted granted Critical
Publication of EP0313826B1 publication Critical patent/EP0313826B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/084—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades the fluid circulating at the periphery of a multistage rotor, e.g. of drum 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the 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
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations

Definitions

  • the present invention relates to an axially flowed through gas turbine with cooling devices for the turbine rotor and its rotor blades according to the preamble of patent claim 1.
  • a gas turbine with cooling thereof allows a higher gas inlet temperature, which increases the efficiency and the performance.
  • the cooling air duct and the cooling air flow and its distribution over the length of the turbine rotor depend on the gas temperatures prevailing in the individual stages of the turbine.
  • the heated cooling air exits into the gas flow.
  • the gas temperature has already dropped so far that the internal cooling of the rotor blades can be dispensed with.
  • the cooling air is taken from the compressor after its last stage and reaches a row of axial bores distributed along the circumference of a flat annular surface of the rotor before the first turbine stage along the outer surface of the section of the shaft or drum located between the compressor and the turbine.
  • the cooling air flow passes through these bores into the cooling ducts of the rotor, at the end of which it, reduced by the portion branched off for cooling the hottest rotor blades, exits into the propellant gas flow and with it into the diffuser.
  • the inflow of the cooling air to the rotor is essentially swirl-free, i.e. without a peripheral component, in the direction of rotation of the drum, it is accelerated on its way to the rotor by the friction on the circumferential surface of the drum in its circumferential direction, albeit in relation to The peripheral speed is not very strong, so that there is still a large difference in speed at the entry into the bores mentioned and into the rotor cooling channels. It must therefore be accelerated to the circumferential rotor speed there. The drum and the rotor must therefore perform pumping work, which moreover increases the cooling air temperature. Like most of the flow through the cooling channels, this represents a loss factor.
  • Another loss is associated with the cooling air flow exiting the blade root of the last stage. It enters the propellant gas flow with a radially, tangentially and axially directed velocity component and forces it radially, so that the hub boundary layer at the diffuser inlet suffers a thickening that is harmful to the recovery.
  • DE-A-34 24 139 proposes that the rotor cooling air should be given a circumferential speed component in the direction of rotation of the rotor after it emerges from the compressor by means of fixed swirl grids with essentially radially directed blades the peripheral speed of the rotor cooling channels, so that the cooling air does not have to be accelerated to them first.
  • the pump work mentioned and the associated losses are thereby eliminated.
  • channels are arranged on the guide blades of the last stage. They are bounded on the one hand by a cover plate provided at the end of the guide vane and on the other hand by a cover band which closes the tip of the blade.
  • Corresponding channels are arranged on the blades of the last stage. They are bounded on the one hand by a cover plate provided on the blade end and on the other hand by the blade root.
  • These channels are designed as vane grids, in which the cooling air expands to the turbine rotor while the work is being carried out. They do nothing to cool the rotor. At the outlet of the last blade ring, the cooling air is immediately transferred to the diffuser.
  • the present invention arose from the task of appropriately guiding both the rotor and blade cooling air and the rotor disk cooling air in their outlet areas at the rotor end into the diffuser to steer that their speed vectors essentially coincide with that of the average exhaust gas flow in the areas mentioned with regard to magnitude and direction.
  • the working capacity of the rotor cooling air is to be largely exploited.
  • the rotor jacket in the area of the last stage is to be cooled more strongly with the same amount of rotor cooling air than is the case with the known constructions.
  • the disc cooling air quantity can thereby be reduced, which reduces the temperature differences within the rotor and thus the thermal stresses in order to achieve an extension of the service life of the turbine rotor.
  • Fig. 1 shows a part of a turbine rotor 1, which is composed of forged rotor disks 2, 3, 4, the along each other on the end faces forged rings are welded.
  • the blades of the rotor blade rings 5 to 9 are inserted in a known manner with their base of double hammer head profile into the correspondingly profiled blade fastening grooves.
  • guide vanes of guide blade rings 11 to 14 are anchored in a guide blade carrier 10 in a manner similar to the rotor blades in the rotor.
  • the guide vane attachments are only indicated schematically.
  • the last stage of the compressor (not shown) is located to the right of the first rotor blade ring 5 of the turbine -
  • the required cooling air flow is removed, whereupon it is given a tangential speed component, which is equal to the peripheral speed of the rotor cooling channels, by a swirl vane grille arranged between the compressor and the first turbine stage, which is described in the aforementioned DE-A-34 24 139.
  • the cooling air then enters the cooling duct system of the turbine at a relative speed of zero in the circumferential direction substantially axially, as indicated by the speed arrow 16, through a series of cooling air bores 15.
  • the cooling air bores 15 which are provided in large numbers distributed over an annular, flat end face 17 in front of the first rotor blade ring, the cooling air passes into an annular groove 18 which widens in cross-section to its circumference, and from this through a series of interrupted annular gaps 19 in front of the first rotor blade ring 5 and between two of the following rotor blade rings as well as through channels 20 in the area of the blade roots finally into blade root channels 21 of the last rotor blade ring 9.
  • the annular gaps 19 are delimited by the circumferential surfaces of the Rotor jacket and by asymmetrical heat accumulation segments 22, 23, which are located between each two blade rings and protect the rotor jacket and the blade roots against overheating by the propellant gas flow.
  • the blade root ducts 20, 21 can expediently be formed from two grooves in the two blades, which adjoin each other in the circumferential direction and adjoin one another in the circumferential direction, and which result in closed ducts. In the case of the almost axially directed blade roots, these channels, as in the case of the blades of the last rotor blade ring 9, can also be provided in the blade grooves themselves.
  • the guide and rotor blades of the most temperature-loaded stages are designed as hollow blades with air cooling.
  • the cooling air is branched off at the blade roots from the cooling air flow described.
  • the elements of the blade cooling are not shown in FIG. 1.
  • the cooling air passes from the blade root channels 21 of the last moving blade ring 9 into a cooling air blade ring 27, which is attached to the rotor body and which has a frusto-conical rotor blade grille 28 just inside its circumference, which, evenly distributed over its circumference, has cooling air blades 31 which are preceded by a rectifier ring 29 which consists of honeycomb-shaped channels 30 distributed over the entire flow cross section.
  • Fig. 2 shows the circled detail II of Fig. 1 on a larger scale and Fig. 3 shows the development of the section III-III shown in Fig. 2 in the form of a conical shell placed through the center of the channel.
  • the rectifier ring 29 has the task of homogenizing the cooling air jets emerging from the blade root channels 21 of the last rotor blades 9 in order to obtain a flow in the channels delimited by the blades 31 that is as free as possible from separation.
  • the cooling air vane ring 27 fulfills part of the object of the invention presented in the introduction by deflecting the flow threads of the cooling air flow in such a way that their speed vectors over the entire circumference of the diffuser hub essentially coincide with the average speed vector of the exhaust gas flow with the loss-reducing effect described at the outset, by the Low-energy boundary layer is supplied with energy at the diffuser hub and its detachment point is shifted downstream. At the same time, the energy of the rotor cooling air is partially used to deliver work to the rotor.
  • the second measure according to the invention which consists in that the cooling air used for cooling the last rotor disk 4 and branched off from the compressor, such as the blade cooling air, also flows out into the diffuser.
  • the disk cooling air passes through two disk air ducts 33 provided in an outer turbine housing base 32 into one that is delimited by the base 32 and an inner turbine housing base 34 disk-shaped cavity 35 is, as indicated by the speed arrows, deflected radially inwards against the rotor axis in this and passes through a series of inner disk air channels 36 provided near the axle in front of the rotor disk 4, where its main part is deflected upwards and via an annular gap 37 and an annular space 38 is blown out through an annular slot 39 into the hub boundary layer.
  • the convexly curved inlet region 40 of the diffuser hub 41 which sucks in the outflowing disk cooling air together with the reactor cooling air through its curvature, also contributes to the intended inflow into the hub boundary layer.
  • the frustoconical surface area 64 of the cooling air vane ring 27 is designed so as to be inclined with respect to the rotor axis and is dimensioned in length in such a way that the exhaust gas flow behind the last moving vane ring 9 is homogenized.
  • a small part of the disk cooling air flowing in through the channel 36 blocks the labyrinth 41 on the end shield.
  • FIG. 4 and 5 show a second embodiment of the rotor cooling air duct.
  • the cooling air enters via a rotor-fixed intermediate channel 44 into a blade grille 45 of a rotor-fixed blade grille ring 46 and out of this into a blade grille 47 of a guide-blade fixed blade grille ring 48, from which it is deflected into end channels 49.
  • the inlet parts of the latter consist of the front half 50 of a vane grille, the profile lugs, in a rotor-fixed vane grille ring 50 ', and the exit region from the rear half 51 of this vane grille in the cooling air vane ring 53.
  • the end channels 49 are shown in FIG.
  • FIG. 6 A further embodiment of the invention is shown in FIG. 6. After the penultimate rotor blade ring 43, the cooling air is guided axially essentially to the end of the rotor blade ring 9 and only then is it blown out into the exhaust gas flow by a cooling air blade ring 63 in the desired direction. After the penultimate rotor blade ring 43, it again passes, as in the embodiment according to FIG.
  • the end channels 61 extend, preferably at an angle to an axis parallel.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (4)

  1. Turbine à gaz à circulation axiale, avec des dispositifs de refroidissement pour le rotor de turbine (1) et ses couronnes d'aubes mobiles (5 - 9), dans laquelle l'air de refroidissement est dérivé du compresseur et, de manière connue en soi, est accéléré en direction tangentielle par un dispositif rotatif de telle sorte que, par rapport à des trous d'air de refroidissement (15) dans le rotor de turbine (1) à travers lesquels l'air de refroidissement pénètre dans le système de guidage de l'air de refroidissement, il présente une vitesse nulle en direction tangentielle, et où il est prévu, pour le guidage de l'air de refroidissement dans la région du dernier étage (9 + 14), des canaux (26, 21, 28; 44, 45, 47, 50, 49, 51, 52, 39; 54, 55, 57, 60, 61, 62), caractérisée en ce que, dans la région de la couronne des aubes directrices (14) du dernier étage, les canaux rectilignes courent dans la surface latérale du rotor et, dans la région de la couronne d'aubes mobiles (9) du dernier étage, ils courent dans le pied des aubes de celle-ci, en ce qu'en aval des canaux rectilignes de la dernière couronne d'aubes mobiles (9) se trouve une grille d'aubes de l'air de refroidissement (28; 51; 62) dans une couronne d'aubes de l'air de refroidissement (27; 53; 63) fixée au rotor de turbine (1), dont les canaux des aubes sont orientés de telle sorte que les vecteurs de vitesse de l'air de refroidissement sortant dans le diffuseur correspondent essentiellement au vecteur de vitesse moyen du courant de gaz d'échappement, en ce que la région d'entrée (40) du moyeu du diffuseur (42) présente un profilage ayant, en coupe axiale, la forme des lignes de courant, et en ce que la surface latérale tronconique (64) de la couronne d'aubes de l'air de refroidissement (27; 53; 63) est inclinée par rapport à l'axe du rotor et est dimensionnée de telle sorte que le courant de gaz d'échappement est uniformisé après la dernière couronne d'aubes mobiles (9).
  2. Turbine à gaz suivant la revendication 1, caractérisée en ce que le canal d'air de refroidissement dans la région de la dernière couronne d'aubes directrices (14) est formé par une rainure annulaire taillée dans le corps de rotor couverte par des segments symétriques d'accumulation de la chaleur (24) et par des passages (26) percés dans les parois (25) de ces segments d'accumulation de la chaleur (24), en ce que pour le guidage de l'air de refroidissement dans la région de la dernière couronne d'aubes mobiles (9) il est prévu des canaux de pied d'aube (21), et en ce qu'une bague redresseuse (29) est disposée dans la couronne d'aubes de l'air de refroidissement (27), avant la grille d'aubes de l'air de refroidissement (28) en regardant dans le sens de l'écoulement.
  3. Turbine à gaz suivant la revendication 1, caractérisée en ce que le guidage de l'air de refroidissement dans la région de la dernière couronne d'aubes directrices (14) se compose de canaux intermédiaires (54) dans la surface latérale du rotor, d'une grille d'aubes (55) fixée sur le rotor à l'extrémité de ces canaux intermédiaires et d'une grille d'aubes (57) dans une couronne de grille d'aubes (58) fixée aux aubes directrices, et en ce que le guidage de l'air de refroidissement dans la région de la dernière couronne d'aubes mobiles (9) comprend une grille d'aubes (60) dans une couronne de grille d'aubes (59) fixée au rotor, laquelle grille d'aube (60) se compose des demi-aubes antérieures formant les nez d'aubes, ensuite des canaux terminaux (61) dans les pieds d'aubes de la dernière couronne d'aubes mobiles (9) ainsi qu'une couronne d'aubes de l'air de refroidissement (63) fixée au rotor avec une grille d'aubes de l'air de refroidissement (62), qui se compose des demi-aubes postérieures.
  4. Turbine à gaz suivant la revendication 1, caractérisée en ce que le guidage de l'air de refroidissement dans la région de la dernière couronne d'aubes directrices (14) comporte des canaux intermédiaires (44) fixés au rotor, une couronne de grille d'aubes (46) fixée au rotor, avec une grille d'aubes (45) incurvée en direction de l'axe du rotor ainsi qu'une grille d'aubes (47) orientée en direction de l'axe du rotor dans une couronne de grille d'aubes (48) fixée aux aubes directrices, et en ce que le guidage de l'air de refroidissement dans la région de la dernière couronne d'aubes mobiles (9) comporte une grille d'aubes (50) dans une couronne de grille d'aubes (50') fixée au rotor, laquelle grille d'aubes (50) se compose des demi-aubes antérieures formant les nez d'aubes, ensuite des canaux terminaux (49) dans la région des pieds d'aubes de la dernière couronne d'aubes mobiles (9) et une couronne d'aubes de l'air de refroidissement (53) fixée au rotor avec une grille d'aubes de l'air de refroidissement (51), qui se compose des demi-aubes postérieures, caractérisée en outre par un espace annulaire (52) et une fente annulaire (39) entre la couronne d'aubes de l'air de refroidissement (53) et le moyeu du diffuseur (42).
EP88115694A 1987-10-30 1988-09-23 Turbine à gaz avec flux axial Expired - Lifetime EP0313826B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE3736836 1987-10-30
DE19873736836 DE3736836A1 (de) 1987-10-30 1987-10-30 Axial durchstroemte gasturbine

Publications (2)

Publication Number Publication Date
EP0313826A1 EP0313826A1 (fr) 1989-05-03
EP0313826B1 true EP0313826B1 (fr) 1992-09-02

Family

ID=6339440

Family Applications (1)

Application Number Title Priority Date Filing Date
EP88115694A Expired - Lifetime EP0313826B1 (fr) 1987-10-30 1988-09-23 Turbine à gaz avec flux axial

Country Status (5)

Country Link
US (1) US4910958A (fr)
EP (1) EP0313826B1 (fr)
JP (1) JP2656576B2 (fr)
CA (1) CA1310273C (fr)
DE (2) DE3736836A1 (fr)

Families Citing this family (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE59102139D1 (de) * 1990-03-23 1994-08-18 Asea Brown Boveri Axialdurchströmte Gasturbine.
DE4324034A1 (de) * 1993-07-17 1995-01-19 Abb Management Ag Gasturbine mit gekühltem Rotor
US6217280B1 (en) 1995-10-07 2001-04-17 Siemens Westinghouse Power Corporation Turbine inter-disk cavity cooling air compressor
DE19617539B4 (de) * 1996-05-02 2006-02-09 Alstom Rotor für eine thermische Turbomaschine
DE19653839A1 (de) * 1996-12-21 1998-06-25 Asea Brown Boveri Rotor eines Turbogenerators mit direkter Gaskühlung
DE19852604A1 (de) * 1998-11-14 2000-05-18 Abb Research Ltd Rotor für eine Gasturbine
DE19854907A1 (de) * 1998-11-27 2000-05-31 Rolls Royce Deutschland Kühlluftführung an einer Axialturbine
DE19854908A1 (de) * 1998-11-27 2000-05-31 Rolls Royce Deutschland Schaufel und Laufscheibe einer Strömungsmaschine
DE19914227B4 (de) * 1999-03-29 2007-05-10 Alstom Wärmeschutzvorrichtung in Gasturbinen
US6402471B1 (en) * 2000-11-03 2002-06-11 General Electric Company Turbine blade for gas turbine engine and method of cooling same
DE102004007327A1 (de) * 2004-02-14 2005-09-15 Alstom Technology Ltd Rotor
GB0503676D0 (en) * 2005-02-23 2005-03-30 Rolls Royce Plc A lock plate arrangement
US8277170B2 (en) 2008-05-16 2012-10-02 General Electric Company Cooling circuit for use in turbine bucket cooling
US8628297B2 (en) 2010-08-20 2014-01-14 General Electric Company Tip flowpath contour
US8591184B2 (en) * 2010-08-20 2013-11-26 General Electric Company Hub flowpath contour
US8784061B2 (en) * 2011-01-31 2014-07-22 General Electric Company Methods and systems for controlling thermal differential in turbine systems
EP2520764A1 (fr) 2011-05-02 2012-11-07 MTU Aero Engines GmbH Aube avec pied refroidi
EP2551453A1 (fr) * 2011-07-26 2013-01-30 Alstom Technology Ltd Dispositif de refroidissement d'un compresseur d'un turbomoteur
US9080449B2 (en) * 2011-08-16 2015-07-14 United Technologies Corporation Gas turbine engine seal assembly having flow-through tube
CH705840A1 (de) 2011-12-06 2013-06-14 Alstom Technology Ltd Hochdruck-Verdichter, insbesondere in einer Gasturbine.
EP2725191B1 (fr) * 2012-10-23 2016-03-16 Alstom Technology Ltd Turbine à gaz et aube de turbine pour une telle turbine à gaz
EP2837769B1 (fr) * 2013-08-13 2016-06-29 Alstom Technology Ltd Arbre de rotor pour turbomachine
US10001061B2 (en) 2014-06-06 2018-06-19 United Technologies Corporation Cooling system for gas turbine engines
EP3106613A1 (fr) * 2015-06-06 2016-12-21 United Technologies Corporation Système de refroidissement pour moteur à turbine à gaz
EP3124742B1 (fr) * 2015-07-28 2018-11-07 MTU Aero Engines GmbH Turbine a gaz
FR3054855B1 (fr) * 2016-08-08 2020-05-01 Safran Aircraft Engines Disque de rotor de turbomachine
DE102022200592A1 (de) 2022-01-20 2023-07-20 Siemens Energy Global GmbH & Co. KG Turbinenschaufel und Rotor
DE102022201077A1 (de) 2022-02-02 2023-08-03 Siemens Energy Global GmbH & Co. KG Verbessertes Nutdesign einer Scheibe für eine Turbinenschaufel, Verfahren und Rotor
DE102022202368A1 (de) 2022-03-10 2023-09-14 Siemens Energy Global GmbH & Co. KG Nutdesign einer Scheibe für eine Turbinenschaufel, Rotor und ein Verfahren

Family Cites Families (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1819864A (en) * 1930-03-24 1931-08-18 Gen Electric Elastic fluid turbine
BE478671A (fr) * 1939-12-19 1948-01-31
US2489683A (en) * 1943-11-19 1949-11-29 Edward A Stalker Turbine
US2713990A (en) * 1948-12-21 1955-07-26 Solar Aircraft Co Exhaust structure for gas turbine
CH340669A (de) * 1956-04-06 1959-08-31 Sulzer Ag Gasturbine mit einem mehrstufigen, mindestens teilweise gekühlten Rotor
GB999611A (en) * 1962-03-07 1965-07-28 Gasturbinenbaw Und Energinmasc Means for cooling turbine discs
CH483557A (de) * 1967-09-12 1969-12-31 Prvni Brnenska Strojirna Zd Y Einrichtung zum Oberflächenschutz von Turbinenläufern, insbesondere von Gasturbinen
IT1063518B (it) * 1975-09-08 1985-02-11 Gen Electric Sistema di utilizzazione della perdita di aria di raffreddamento in un turbomotore a gas
CH594809A5 (fr) * 1975-10-10 1978-01-31 Bbc Brown Boveri & Cie
GB1524956A (en) * 1975-10-30 1978-09-13 Rolls Royce Gas tubine engine
US4186554A (en) * 1975-11-10 1980-02-05 Possell Clarence R Power producing constant speed turbine
US4113406A (en) * 1976-11-17 1978-09-12 Westinghouse Electric Corp. Cooling system for a gas turbine engine
GB2081392B (en) * 1980-08-06 1983-09-21 Rolls Royce Turbomachine seal
US4456427A (en) * 1981-06-11 1984-06-26 General Electric Company Cooling air injector for turbine blades
GB2118629B (en) * 1982-04-21 1985-07-17 Rolls Royce Device for passing a fluid flow eg. cooling air through a barrier eg. bolted joint
DE3424139C2 (de) * 1984-06-30 1996-02-22 Bbc Brown Boveri & Cie Gasturbinenrotor
US4666368A (en) * 1986-05-01 1987-05-19 General Electric Company Swirl nozzle for a cooling system in gas turbine engines

Also Published As

Publication number Publication date
DE3874283D1 (de) 1992-10-08
JPH01151725A (ja) 1989-06-14
US4910958A (en) 1990-03-27
JP2656576B2 (ja) 1997-09-24
DE3736836A1 (de) 1989-05-11
EP0313826A1 (fr) 1989-05-03
CA1310273C (fr) 1992-11-17

Similar Documents

Publication Publication Date Title
EP0313826B1 (fr) Turbine à gaz avec flux axial
EP0447886B1 (fr) Turbine à gaz avec flux axiale
DE60318792T2 (de) Zapfluft-Gehäuse für einen Verdichter
DE60214137T2 (de) Verfahren und vorrichtung zur kühlung der schaufelspitzen in einer gasturbine
DE69926574T2 (de) Turbinenleitgitter mit einem Kühlluftleitsystem
DE60132405T2 (de) Kühlung von Turbinenschaufeln durch spezifische Schaufelverteilung
EP1736635B1 (fr) Système de transfert d'air entre le compresseur et la turbine d'une turbine à gaz
DE3713923C2 (de) Kühlluft-Übertragungsvorrichtung
DE2943464A1 (de) Dichtungsvorrichtung fuer ein gasturbinentriebwerk
EP2179143B1 (fr) Refroidissement de fente entre une paroi de chambre de combustion et une paroi de turbine d'une installation de turbine à gaz
DE1130646B (de) Diagonalgasturbinen-Energieanlage
EP3093447B1 (fr) Rotor d'une turbine a gaz ayant un guidage d'air de refroidissement ameliore
DE2261443A1 (de) Turbinenanordnung mit zweistromkuehlung fuer gasturbinentriebwerke
EP1111189B1 (fr) Chemin d'air de refroidissement pour le rotor d'une turbine à gaz
DE1601564A1 (de) Mantelring fuer Gasturbinenanlagen
DE2147537A1 (de) Kühleinrichtung für die Enden von Turbinenlaufschaufeln mit Luftexpansion
DE665762C (de) Einrichtung zur Kuehlung von Turbinen, insbesondere Gasturbinen
DE2624312A1 (de) Turbine, insbesondere fuer einen turbolader
DE2356721B2 (de) Kühleinrichtung für hohle Laufschaufeln einer axial durchströmten Turbine
DE2507182A1 (de) Axialgasturbinenanlage
DE102009040758A1 (de) Umlenkvorrichtung für einen Leckagestrom in einer Gasturbine und Gasturbine
EP1656497B1 (fr) Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz
EP3336313A1 (fr) Ensemble d'aube mobile pour turbines d'une turbine turbine à gaz et procédé de fourniture d'air sceau dans un ensemble d'aube mobile pour turbines
EP0992656B1 (fr) Turbomachine pour comprimer ou détendre un fluide comprimable
CH663251A5 (de) Einrichtung zur kuehlung der rotoren von dampfturbinen.

Legal Events

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

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): CH DE GB IT LI NL SE

17P Request for examination filed

Effective date: 19891028

17Q First examination report despatched

Effective date: 19901012

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): CH DE GB IT LI NL SE

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: LI

Effective date: 19920930

Ref country code: CH

Effective date: 19920930

REF Corresponds to:

Ref document number: 3874283

Country of ref document: DE

Date of ref document: 19921008

ITF It: translation for a ep patent filed

Owner name: DE DOMINICIS & MAYER S.

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)
REG Reference to a national code

Ref country code: CH

Ref legal event code: PL

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

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

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: SE

Payment date: 19930817

Year of fee payment: 6

26N No opposition filed
PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Effective date: 19940924

EAL Se: european patent in force in sweden

Ref document number: 88115694.7

EUG Se: european patent has lapsed

Ref document number: 88115694.7

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 19970811

Year of fee payment: 10

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: NL

Payment date: 19970821

Year of fee payment: 10

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 19970924

Year of fee payment: 10

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 19980923

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: NL

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 19990401

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 19980923

NLV4 Nl: lapsed or anulled due to non-payment of the annual fee

Effective date: 19990401

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 19990701

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES;WARNING: LAPSES OF ITALIAN PATENTS WITH EFFECTIVE DATE BEFORE 2007 MAY HAVE OCCURRED AT ANY TIME BEFORE 2007. THE CORRECT EFFECTIVE DATE MAY BE DIFFERENT FROM THE ONE RECORDED.

Effective date: 20050923