EP2861829A1 - Aube mobile de turbine à gaz à refroidissement par air - Google Patents

Aube mobile de turbine à gaz à refroidissement par air

Info

Publication number
EP2861829A1
EP2861829A1 EP13729666.1A EP13729666A EP2861829A1 EP 2861829 A1 EP2861829 A1 EP 2861829A1 EP 13729666 A EP13729666 A EP 13729666A EP 2861829 A1 EP2861829 A1 EP 2861829A1
Authority
EP
European Patent Office
Prior art keywords
blade
coolant
turbine
span
airfoil
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Withdrawn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
EP13729666.1A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Nihal Kurt
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
Publication of EP2861829A1 publication Critical patent/EP2861829A1/fr
Withdrawn legal-status Critical Current

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
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting

Definitions

  • the invention relates to an air-cooled turbine blade for a gas turbine having a blade root and an adjoining aerodynamically-curved airfoil having a suction-side sidewall and a pressure-side sidewall extending in a chordal direction from a leading edge to a trailing edge and in a span span with a total span extend from a blade end to a blade tip end, wherein a coolant channel is provided in the blade for guiding a coolant.
  • Such turbine blades are well known in the widely available prior art.
  • the turbine rotor blades which are generally produced by casting, are hollow with the aid of casting cores used in the casting apparatus, so that cooling channels are present in the interior through which a coolant, usually cooling air, can flow during operation.
  • the coolant provides sufficient cooling of the turbine blade material to achieve the predetermined and expected life, despite the high ambient temperatures encountered in the operation of a gas turbine in the vicinity of the turbine blade.
  • the cooling channels are now diverted several times in meandering form.
  • turbulators are often provided on the inner sides of the blade outer walls in order to increase the heat transfer from the material of the turbine blade into the cooling air.
  • the then heated cooling air will either be at the trailing edge of the airfoil, at the top of the blade plate or in the region of the leading edge of the blade by corresponding openings, which may possibly also be configured as a film cooling holes, blown out and mixed in the hot gas flow.
  • the design of a turbine blade is particularly difficult, especially at lower profile thickness.
  • the wall thicknesses of the airfoil and also the cooling channels provided in the interior must be made comparatively filigree in order to be able to meet the requirement for a low profile thickness.
  • Lower cooling channel cross-sections lead to lower GusskernwandCn so that they tend to break when handling and the assembly of the casting device.
  • the turbine blade according to the invention for a gas turbine with a blade root and an adjoining aerodynamically curved blade having a suction-side side wall and a pressure-side side wall in a chord direction from a common leading edge to a trailing edge and in a Spannweiterich- direction a total span ranging from a blade end to a blade tip end, wherein in the blade for guiding the coolant, a coolant channel is provided, it is provided that the blade, except going from its blade end equal to 0% of the blade span from a span of 60% of the total span, preferably from 75% of the total span, is coolant channel free.
  • this is designed integrally and thus produced in the casting process.
  • the invention is based on the finding that such turbine blades do not have to withstand the highest currently possible hot gas temperatures, but rather lower temperatures, as can occur, for example, in large, stationary gas turbines in the second or third turbine stage.
  • the special feature of the turbine blade according to the invention is that it can be designed massive pointed, so that only the central region of the blade profile in Spannweiteterrorism and the foot-side portion of the airfoil is to cool.
  • the invention takes into account the knowledge that viewed in Spannweiteiques, the hottest temperatures occur in the central region of the span, whereas in the outer edge portions - ie shovel tip side and Schaufel foot side - lower temperatures. In this respect, a blade tip-side cooling of the blade is not required, so this can be coolant channel free.
  • the coolant channel free area makes it possible to produce turbine blades with a comparatively small profile thickness, as this range increases the overall rigidity and strength of the airfoil.
  • the coolant channel has an inlet arranged in the blade root for coolant and at least one coolant outlet, which coolant outlet or coolant outlet is also arranged exclusively in the blade root.
  • Airfoil heating-up cooling air is not introduced via the airfoil directly into the hot gas flow in the hot gas path of the turbine, but blown out in a region outside the hot gas path of the gas turbine.
  • the temperature gradient of the components limiting the hot gas path can be reduced, since their colder sides can be tempered with the help of the blown, but preheated cooling air.
  • the coolant outlet may even be provided on the inflow side, but nevertheless
  • FIG. 1 shows a turbine blade in a longitudinal section.
  • Turbine blade 10 is for a medium or low cooled turbine stage of a stationary gas turbine. As such, the turbine bucket 10 may be used in a second, third, or fourth turbine stage.
  • the turbine blade 10 has a blade root 12 shown in FIG. 1 below.
  • the blade root 12 in this case comprises a platform 14, in which in span direction an airfoil 16 extends.
  • the blade airfoil 16 extends in the chordwise direction from a leading edge 18 to a trailing edge 20.
  • the airfoil 16 is surrounded by a hot gas during operation in a turbine blade 10 used in a gas turbine, so that the leading edge 18 is located downstream and the trailing edge 20 downstream.
  • the airfoil 20 has a blade end-side end 22 and a blade tip-side end 24.
  • the average total span is determined by the arithmetic mean of the span at the leading edge 18 and the span at the trailing edge 20 and is normalized to a value of 100%.
  • the origin of the span is located in the transition from platform 14 to the blade root end 22 of the airfoil 16, so that the overall span of 100% is blade-tipped.
  • the center of the span at 50% of the airfoil is designated by reference numeral 26.
  • the turbine bucket 10 is hollow with a single coolant channel 28 having a blade-side inlet 30. Downstream of the single deflection section 32, the coolant channel 28 comprises a leading edge section 34, the outlet 36 of which is arranged in FIG.
  • the turbine blade 10 can have sufficient stability and strength despite a comparatively small profile thickness. At the same time, it is cooled sufficiently to achieve the required life.
  • the internally disposed coolant channel 28 may also be shaped differently or also have a downstream outlet 36.
  • An airfoil 16 having a suction side sidewall and a pressure sidewall extending in a chordwise direction from a common leading edge 18 to a trailing edge 20 and in a spanwise direction with a total span from a blade root end 22 to a blade tip end 24, wherein in the airfoil 16 to guide a coolant, a coolant channel 28 is provided.
  • a sufficiently coolable turbine blade 10 which may be provided with a comparatively small profile thickness, it is proposed that the blade 16, starting from its base foot end 22 from a span of 75% of the total span, preferably 60% of the total span, cool - is medium-channel-free.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne une aube mobile (10 ) de turbine à gaz à refroidissement par air, comportant un pied (12) et une aube (16) aérodynamiquement bombée et reliée audit pied, ladite aube présentant une paroi latérale côté aspiration et une paroi latérale côté pression qui s'étendent dans une direction de la corde, d'un bord d'attaque commun (18) à un bord de fuite (20), et dans une direction de l'envergure, avec une envergure commune, d'une extrémité côté pied d'aube (22) à une extrémité côté pointe d'aube (24). Un conduit de liquide de refroidissement (28) situé dans l'aube (16) sert à guider un liquide de refroidissement. Selon l'invention, afin d'obtenir une aube de turbine (10) qui peut être suffisamment refroidie et qui présente une épaisseur de profil relativement faible, l'aube (16) est exempte de conduit de fluide de refroidissement en partant de son extrémité côté pied d'aube (22) à partir d'une distance correspondant à 75% de l'envergure totale, de préférence à 60% de l'envergure totale. FIG
EP13729666.1A 2012-07-12 2013-06-11 Aube mobile de turbine à gaz à refroidissement par air Withdrawn EP2861829A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102012212235.4A DE102012212235A1 (de) 2012-07-12 2012-07-12 Turbinenlaufschaufel für eine Gasturbine
PCT/EP2013/061952 WO2014009075A1 (fr) 2012-07-12 2013-06-11 Aube mobile de turbine à gaz à refroidissement par air

Publications (1)

Publication Number Publication Date
EP2861829A1 true EP2861829A1 (fr) 2015-04-22

Family

ID=48652030

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13729666.1A Withdrawn EP2861829A1 (fr) 2012-07-12 2013-06-11 Aube mobile de turbine à gaz à refroidissement par air

Country Status (7)

Country Link
US (1) US20150167476A1 (fr)
EP (1) EP2861829A1 (fr)
JP (1) JP2015522127A (fr)
CN (1) CN104471191A (fr)
DE (1) DE102012212235A1 (fr)
RU (1) RU2015104667A (fr)
WO (1) WO2014009075A1 (fr)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180066525A1 (en) * 2016-09-02 2018-03-08 James P. Downs Air cooled turbine rotor blade for closed loop cooling
FR3081912B1 (fr) * 2018-05-29 2020-09-04 Safran Aircraft Engines Aube de turbomachine comprenant un passage interne d'ecoulement de fluide equipe d'une pluralite d'elements perturbateurs a agencement optimise

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE559676C (de) * 1931-08-20 1932-09-22 E H Hans Holzwarth Dr Ing Verfahren zur Kuehlung von Schaufeln, insbesondere fuer Brennkraftturbinen
DE876936C (de) * 1948-10-01 1953-04-02 Aachen Dr.-Ing. Karl Leist Arbeitsverfahren ifür Gasturbinen und Gasturbine zur Ausübung des Verfahrens
GB677052A (en) * 1949-12-28 1952-08-06 Bbc Brown Boveri & Cie Cooled turbine rotor
FR2275975A5 (fr) * 1973-03-20 1976-01-16 Snecma Perfectionnements au refroidissement d'aubes de turbines a gaz
US4645424A (en) * 1984-07-23 1987-02-24 United Technologies Corporation Rotating seal for gas turbine engine
KR100389990B1 (ko) * 1995-04-06 2003-11-17 가부시끼가이샤 히다치 세이사꾸쇼 가스터빈
JP3621523B2 (ja) * 1996-09-25 2005-02-16 株式会社東芝 ガスタービンの動翼冷却装置
JP3475838B2 (ja) * 1999-02-23 2003-12-10 株式会社日立製作所 タービンロータ及びタービンロータのタービン動翼冷却方法
US7137782B2 (en) * 2004-04-27 2006-11-21 General Electric Company Turbulator on the underside of a turbine blade tip turn and related method

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
None *
See also references of WO2014009075A1 *

Also Published As

Publication number Publication date
RU2015104667A (ru) 2016-08-27
WO2014009075A1 (fr) 2014-01-16
CN104471191A (zh) 2015-03-25
US20150167476A1 (en) 2015-06-18
JP2015522127A (ja) 2015-08-03
DE102012212235A1 (de) 2014-01-16

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