EP0313826A1 - Turbine à gaz avec flux axial - Google Patents

Turbine à gaz avec flux axial Download PDF

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Publication number
EP0313826A1
EP0313826A1 EP88115694A EP88115694A EP0313826A1 EP 0313826 A1 EP0313826 A1 EP 0313826A1 EP 88115694 A EP88115694 A EP 88115694A EP 88115694 A EP88115694 A EP 88115694A EP 0313826 A1 EP0313826 A1 EP 0313826A1
Authority
EP
European Patent Office
Prior art keywords
rotor
blade
cooling air
ring
last
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
EP88115694A
Other languages
German (de)
English (en)
Other versions
EP0313826B1 (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

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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
    • 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 flow-through gas turbine with cooling devices for the turbine rotor and its rotor blades, the cooling air being branched off from the compressor and accelerated in a known manner by a swirl device in the circumferential direction so that it is opposite cooling air holes on the turbine rotor, through which the cooling air into the Cooling air system flows in, has zero speed in the circumferential direction.
  • 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. You only get cooling in the area of the blade roots through the air flowing towards the end of the rotor body, which exits into the already largely relaxed propellant gas stream before and after the foot area of the last row of blades and reaches the exhaust gas diffuser with it.
  • 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 channels 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. 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 away, so that the hub boundary layer at the diffuser inlet suffers a thickening that is harmful to the recovery.
  • the present invention arose from the task of guiding the rotor and blade cooling air as well as the rotor disk cooling air in their outlet areas at the rotor end into the diffuser in such a way that their velocity vectors correspond to that of the average exhaust gas flow in the areas mentioned in terms of amount and direction essentially coincide.
  • the working capacity of the rotor cooling air should be largely used.
  • This guide is also intended to cool the rotor jacket in the area of the last stage 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.
  • the axially flowed through gas turbine is characterized in that for the cooling air duct in the In the area of the last stage, channels are provided which run in the area of the guide vane ring of the last stage in the rotor casing and in the area of the rotor vane ring of the last stage in its blade roots, a cooling air vane grille being present in a cooling vane ring attached to the turbine rotor, at least at the end of the last rotor vane ring Channels are oriented so that the speed vectors of the cooling air exiting into the diffuser essentially coincide with the average speed vector of the exhaust gas flow, and the limits for the outflow of the cooling air into the diffuser are designed in such a way that their separation is avoided and the propellant gas flow in the hub area of the last blade ring is homogenized.
  • Fig. 1 shows a part of a turbine rotor 1, which is composed of forged rotor disks 2, 3, 4, which along with each other on the end faces forged rings who 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 two rotor blade rings and protect the rotor jacket and the rotor blade feet from overheating by the propellant gas flow.
  • the blade root passages 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 passages. In the case of the almost axially directed blade roots, these channels can also be provided in the blade grooves themselves, as in the blades of the last rotor blade ring 9.
  • 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 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, flows out into the diffuser in a guided manner.
  • the disk cooling air passes through two disk air channels 33 provided in an outer turbine housing base 32 into one bounded by the bottom 32 and an inner turbine housing base 34 th 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 upward 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 area 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 lateral surface 64 of the cooling air vane ring 27 is designed to be inclined with respect to the rotor axis and its length is such that the exhaust gas flow behind the last rotor 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 Another 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 is only then blown out into the exhaust gas flow in the desired direction by a cooling air blade ring 63. After the penultimate rotor blade ring 43, it again passes, as in the embodiment according to FIG.
  • the end channels 61 extend between the two vane grids 60 and 61, preferably inclined 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)
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 true EP0313826A1 (fr) 1989-05-03
EP0313826B1 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)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0447886A1 (fr) * 1990-03-23 1991-09-25 Asea Brown Boveri Ag Turbine à gaz avec flux axiale
EP0636764A1 (fr) * 1993-07-17 1995-02-01 ABB Management AG Turbine à gaz avec refroidissement du rotor
WO1999047798A1 (fr) * 1998-03-16 1999-09-23 Siemens Westinghouse Power Corporation Mecanisme changeur de pression destine a un air de refroidissement de turbine
US8277170B2 (en) 2008-05-16 2012-10-02 General Electric Company Cooling circuit for use in turbine bucket cooling
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
EP3106613A1 (fr) * 2015-06-06 2016-12-21 United Technologies Corporation Système de refroidissement pour moteur à turbine à gaz
FR3054855A1 (fr) * 2016-08-08 2018-02-09 Safran Aircraft Engines Disque de rotor de turbomachine
US10001061B2 (en) 2014-06-06 2018-06-19 United Technologies Corporation Cooling system for gas turbine engines

Families Citing this family (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
DE19854908A1 (de) * 1998-11-27 2000-05-31 Rolls Royce Deutschland Schaufel und Laufscheibe einer Strömungsmaschine
DE19854907A1 (de) * 1998-11-27 2000-05-31 Rolls Royce Deutschland Kühlluftführung an einer Axialturbine
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
US8591184B2 (en) * 2010-08-20 2013-11-26 General Electric Company Hub flowpath contour
US8628297B2 (en) 2010-08-20 2014-01-14 General Electric Company Tip flowpath contour
US8784061B2 (en) * 2011-01-31 2014-07-22 General Electric Company Methods and systems for controlling thermal differential in turbine systems
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
EP3124742B1 (fr) * 2015-07-28 2018-11-07 MTU Aero Engines GmbH Turbine a gaz
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

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US1819864A (en) * 1930-03-24 1931-08-18 Gen Electric Elastic fluid turbine
US2489683A (en) * 1943-11-19 1949-11-29 Edward A Stalker Turbine
CH270345A (de) * 1939-12-19 1950-08-31 Power Jets Res & Dev Ltd Gasturbinen-Kraftanlage.
US2713990A (en) * 1948-12-21 1955-07-26 Solar Aircraft Co Exhaust structure for gas turbine
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
DE2549112A1 (de) * 1975-10-10 1977-04-21 Bbc Brown Boveri & Cie Turbinenkuehlung

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CH340669A (de) * 1956-04-06 1959-08-31 Sulzer Ag Gasturbine mit einem mehrstufigen, mindestens teilweise gekühlten Rotor
IT1063518B (it) * 1975-09-08 1985-02-11 Gen Electric Sistema di utilizzazione della perdita di aria di raffreddamento in un turbomotore a gas
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

Patent Citations (7)

* 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
CH270345A (de) * 1939-12-19 1950-08-31 Power Jets Res & Dev Ltd Gasturbinen-Kraftanlage.
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
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
DE2549112A1 (de) * 1975-10-10 1977-04-21 Bbc Brown Boveri & Cie Turbinenkuehlung

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0447886A1 (fr) * 1990-03-23 1991-09-25 Asea Brown Boveri Ag Turbine à gaz avec flux axiale
US5189874A (en) * 1990-03-23 1993-03-02 Asea Brown Boveri Ltd. Axial-flow gas turbine cooling arrangement
EP0636764A1 (fr) * 1993-07-17 1995-02-01 ABB Management AG Turbine à gaz avec refroidissement du rotor
US6217280B1 (en) 1995-10-07 2001-04-17 Siemens Westinghouse Power Corporation Turbine inter-disk cavity cooling air compressor
WO1999047798A1 (fr) * 1998-03-16 1999-09-23 Siemens Westinghouse Power Corporation Mecanisme changeur de pression destine a un air de refroidissement de turbine
US8277170B2 (en) 2008-05-16 2012-10-02 General Electric Company Cooling circuit for use in turbine bucket cooling
EP2520764A1 (fr) * 2011-05-02 2012-11-07 MTU Aero Engines GmbH Aube avec pied refroidi
US9739151B2 (en) 2011-05-02 2017-08-22 Mtu Aero Engines Gmbh Blade, integrally bladed rotor base body and turbomachine
US9382802B2 (en) 2011-07-26 2016-07-05 General Electric Technology Gmbh Compressor rotor
EP2551453A1 (fr) * 2011-07-26 2013-01-30 Alstom Technology Ltd Dispositif de refroidissement d'un compresseur d'un turbomoteur
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
FR3054855A1 (fr) * 2016-08-08 2018-02-09 Safran Aircraft Engines Disque de rotor de turbomachine
WO2018029408A1 (fr) * 2016-08-08 2018-02-15 Safran Aircraft Engines Disque de rotor de turbomachine
GB2567103A (en) * 2016-08-08 2019-04-03 Safran Aircraft Engines Turbo engine rotor disc
US10954795B2 (en) 2016-08-08 2021-03-23 Safran Aircraft Engines Turbo engine rotor disc
GB2567103B (en) * 2016-08-08 2022-01-26 Safran Aircraft Engines Turbo engine rotor disc

Also Published As

Publication number Publication date
EP0313826B1 (fr) 1992-09-02
JP2656576B2 (ja) 1997-09-24
DE3736836A1 (de) 1989-05-11
JPH01151725A (ja) 1989-06-14
US4910958A (en) 1990-03-27
DE3874283D1 (de) 1992-10-08
CA1310273C (fr) 1992-11-17

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