EP0892151A1 - Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas - Google Patents

Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas Download PDF

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Publication number
EP0892151A1
EP0892151A1 EP97810492A EP97810492A EP0892151A1 EP 0892151 A1 EP0892151 A1 EP 0892151A1 EP 97810492 A EP97810492 A EP 97810492A EP 97810492 A EP97810492 A EP 97810492A EP 0892151 A1 EP0892151 A1 EP 0892151A1
Authority
EP
European Patent Office
Prior art keywords
blade
channel
leading edge
cooling system
pressure
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Withdrawn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
EP97810492A
Other languages
German (de)
English (en)
Inventor
Bernhard Dr. Weigand
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.)
General Electric Switzerland GmbH
Original Assignee
ABB Asea Brown Boveri Ltd
Asea Brown Boveri AB
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 ABB Asea Brown Boveri Ltd, Asea Brown Boveri AB filed Critical ABB Asea Brown Boveri Ltd
Priority to EP97810492A priority Critical patent/EP0892151A1/fr
Priority to US09/111,874 priority patent/US6168380B1/en
Priority to CN98116043A priority patent/CN1113153C/zh
Priority to JP10200528A priority patent/JPH1172005A/ja
Publication of EP0892151A1 publication Critical patent/EP0892151A1/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Withdrawn legal-status Critical Current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
    • 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

Definitions

  • the invention relates to a cooling system for the leading edge area of a hollow Gas turbine blade, in which a channel is located within the thickened blade leading edge extends from the blade root to the blade tip, the channel over a plurality of bores made in the front edge of the blade communicates with a main channel through which a coolant flows.
  • Cooling poses a particular problem of the leading edge area of such blades.
  • a cooling system of the type mentioned is known from DE-A1 27 03 815.
  • the blade there has a main channel in the leading edge area, the formed by an insert supported on the inner walls of the blade becomes.
  • the front edge section is thickened and encloses one Cavity.
  • the thickened section is both with the blade root and with connected to the blade cover plate and is used in particular for torsional stiffening.
  • the height of the cavity is increased by several holes with coolant fed from the main canal with longitudinal flow.
  • the inside pages the front edge in the area of the cavity is chilled.
  • the cavity is Provide through holes on the actual front edge to the outer wall.
  • the coolant exiting through the through holes in the turbine duct thus causes film cooling of the leading edge area.
  • the holes from The main channel to the cavity is dimensioned in such a way that the one for the subsequent one Film cooling required pressure drop is caused.
  • the present invention has for its object a cooling system of the beginning to create mentioned type, in which the front edge with pure convection cooling is applied without additional film cooling.
  • the cast blade shown in Fig. 1 has three inner chambers a, b, and c by a coolant, such as air, perpendicular to the plane of the drawing are flowed through.
  • a coolant such as air
  • the inside of the blade contour forming wall W - which is surrounded on both sides by hot gases - by the Coolant flows around and give off their heat to the coolant.
  • Tools such as guide ribs, flow channels, inserts for impingement cooling and the like can be provided to improve wall cooling.
  • a blade provided with a cover plate 11 circulates this Coolant in several passes through the inner chambers a, b, and c and can e.g. are discharged into the turbine duct via the trailing edge of the blade, not shown.
  • FIG. 2 to 5 show the cooling system for the front edge area of a hollow Gas turbine blade. It extends from the blade root 1 to the blade tip 2 a longitudinally flowed main channel 3, which corresponds to chamber a in FIG. 1. In the area of the airfoil 4, this channel is from the inner walls of the Front edge, the suction side 6 and the pressure side 7 and one of the pressure side limited with web 8 connecting the suction side.
  • a channel 10 extends within the thickened blade leading edge 5 from Blade root to the tip of the blade. It is understood that this channel depends on Requirements do not have to extend to the blade root. Its lower end could also be a little further radially outwards and e.g. only below middle Use the sheet height where the greatest heat load of the Shovel occurs.
  • the channel 10 merges into one below the cover plate 11 extending chamber 12.
  • This chamber extends to the not shown Blade trailing edge, at least in the chamber area against the flow Gas turbine duct is open.
  • the one prevailing on the trailing edge of the blade Pressure that is in any case lower than that in the main duct 3 through which the longitudinal flow flows prevailing pressure, is therefore effective in channel 10. This pressure difference leads to the fact that the medium located in channel 10 against the rear edge flows out.
  • the trailing edge pressure is not absolutely necessary for this driving pressure difference must be applied to channel 10. So could the chamber 12 are also operatively connected to a vortex chamber, as is usually the case in the labyrinths above the cover plate between two cover plate serrations or Sealing strips are provided.
  • the channel 10 communicates via a plurality of in the interior of the Blade front edge drilled holes 9 with the coolant main duct with longitudinal flow 3.
  • the driving pressure difference ensures that a part of the medium flowing along the front edge in the main channel 3 now flows through these bores 9 into the channel 10 and there as an impact jet strikes the inner wall of the duct.
  • Increased radial extent So more and more cooling air in the duct 10.
  • the channel is expanded in the radial direction.
  • the cross section through which the flow is Blade root up to the tip of the blade increasingly larger, depending on the type from the newly added impact beams.
  • the number and dimensioning of the bores 9 can increase the cross section thus be either steady or discontinuous. Decisive for the type of cross-sectional increase is the requirement that in any case the speed ratio of the respective impact jet to the speed of the longitudinal flow in channel 10 should be big. This prevents the outflowing air from having the effect of Impact rays impaired.
  • This increased heat transfer coefficient applies to the actual nose is convectively cooled by longitudinal and impingement flow.
  • An increased heat transfer coefficient is also in the rear area of the front edge achieved in that the outflow from the channel 3 into the bores 9 Flow intensity increased in this area. Opposite the smooth triangular channel a Without the new measure, considerably more coolant flows along with the along the channel wall provided with holes with correspondingly more effective Cooling.
  • the shape of which e.g. corresponds to the shape of the blade profile the inside wall of the cover plate must be tipped. With this measure, the outflowing air also contribute to cooling the cover plate.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP97810492A 1997-07-15 1997-07-15 Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas Withdrawn EP0892151A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP97810492A EP0892151A1 (fr) 1997-07-15 1997-07-15 Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas
US09/111,874 US6168380B1 (en) 1997-07-15 1998-07-08 Cooling system for the leading-edge region of a hollow gas-turbine blade
CN98116043A CN1113153C (zh) 1997-07-15 1998-07-14 燃气轮机空心叶片前缘区的冷却系统
JP10200528A JPH1172005A (ja) 1997-07-15 1998-07-15 中空のガスタービン羽根の前方縁部領域のための冷却機構

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP97810492A EP0892151A1 (fr) 1997-07-15 1997-07-15 Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas

Publications (1)

Publication Number Publication Date
EP0892151A1 true EP0892151A1 (fr) 1999-01-20

Family

ID=8230305

Family Applications (1)

Application Number Title Priority Date Filing Date
EP97810492A Withdrawn EP0892151A1 (fr) 1997-07-15 1997-07-15 Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas

Country Status (4)

Country Link
US (1) US6168380B1 (fr)
EP (1) EP0892151A1 (fr)
JP (1) JPH1172005A (fr)
CN (1) CN1113153C (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1154124A1 (fr) * 2000-05-10 2001-11-14 General Electric Company Aube refroidie par impact
EP1201879A3 (fr) * 2000-10-27 2003-07-16 ALSTOM (Switzerland) Ltd Composant refroidi, noyau de coulage et procédé pour la fabrication dudit composant
DE102007008319A1 (de) 2007-02-16 2008-08-21 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Prallluftkühlung für Gasturbinen

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7665968B2 (en) * 2004-05-27 2010-02-23 United Technologies Corporation Cooled rotor blade
US8083485B2 (en) * 2007-08-15 2011-12-27 United Technologies Corporation Angled tripped airfoil peanut cavity
US8397516B2 (en) * 2009-10-01 2013-03-19 General Electric Company Apparatus and method for removing heat from a gas turbine
CN102146810A (zh) * 2010-02-10 2011-08-10 中国科学院工程热物理研究所 利用工质的超临界特性对高温涡轮叶片进行冷却的方法
US10041743B2 (en) 2013-01-07 2018-08-07 Carrier Corporation Energy recovery ventilator
JP6245740B2 (ja) * 2013-11-20 2017-12-13 三菱日立パワーシステムズ株式会社 ガスタービン翼
EP3000970B1 (fr) 2014-09-26 2019-06-12 Ansaldo Energia Switzerland AG Système de refroidissement pour le bord d'attaque d'une aube de turbine d'une turbine à gaz
US10077667B2 (en) * 2015-05-08 2018-09-18 United Technologies Corporation Turbine airfoil film cooling holes
US10508554B2 (en) 2015-10-27 2019-12-17 General Electric Company Turbine bucket having outlet path in shroud
US9885243B2 (en) 2015-10-27 2018-02-06 General Electric Company Turbine bucket having outlet path in shroud
US10156145B2 (en) * 2015-10-27 2018-12-18 General Electric Company Turbine bucket having cooling passageway

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1986002406A1 (fr) * 1984-10-10 1986-04-24 Paul Marius A Moteur a turbine a gaz
US4820123A (en) * 1988-04-25 1989-04-11 United Technologies Corporation Dirt removal means for air cooled blades
US5122033A (en) * 1990-11-16 1992-06-16 Paul Marius A Turbine blade unit

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1565361A (en) 1976-01-29 1980-04-16 Rolls Royce Blade or vane for a gas turbine engien
US4514144A (en) * 1983-06-20 1985-04-30 General Electric Company Angled turbulence promoter
US4820122A (en) * 1988-04-25 1989-04-11 United Technologies Corporation Dirt removal means for air cooled blades
US5403159A (en) * 1992-11-30 1995-04-04 United Technoligies Corporation Coolable airfoil structure

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1986002406A1 (fr) * 1984-10-10 1986-04-24 Paul Marius A Moteur a turbine a gaz
US4820123A (en) * 1988-04-25 1989-04-11 United Technologies Corporation Dirt removal means for air cooled blades
US5122033A (en) * 1990-11-16 1992-06-16 Paul Marius A Turbine blade unit

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1154124A1 (fr) * 2000-05-10 2001-11-14 General Electric Company Aube refroidie par impact
US6435813B1 (en) 2000-05-10 2002-08-20 General Electric Company Impigement cooled airfoil
EP1201879A3 (fr) * 2000-10-27 2003-07-16 ALSTOM (Switzerland) Ltd Composant refroidi, noyau de coulage et procédé pour la fabrication dudit composant
DE102007008319A1 (de) 2007-02-16 2008-08-21 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Prallluftkühlung für Gasturbinen
US8152463B2 (en) 2007-02-16 2012-04-10 Rolls-Royce Deutschland Ltd & Co Kg Method for impingement air cooling for gas turbines

Also Published As

Publication number Publication date
CN1205389A (zh) 1999-01-20
US6168380B1 (en) 2001-01-02
JPH1172005A (ja) 1999-03-16
CN1113153C (zh) 2003-07-02

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