EP0475658A1 - Turbinenschaufel mit serieller Stossstrahlkühlung durch interne kammerformende Rippen - Google Patents

Turbinenschaufel mit serieller Stossstrahlkühlung durch interne kammerformende Rippen Download PDF

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Publication number
EP0475658A1
EP0475658A1 EP91308047A EP91308047A EP0475658A1 EP 0475658 A1 EP0475658 A1 EP 0475658A1 EP 91308047 A EP91308047 A EP 91308047A EP 91308047 A EP91308047 A EP 91308047A EP 0475658 A1 EP0475658 A1 EP 0475658A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
cavities
series
air flow
leading
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Withdrawn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
EP91308047A
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English (en)
French (fr)
Inventor
Ching-Pang Lee
Chung-Der Young
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 Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP0475658A1 publication Critical patent/EP0475658A1/de
Withdrawn legal-status Critical Current

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    • 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
    • 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
    • F01D5/186—Film 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence

Definitions

  • the present invention relates generally to gas turbine engine blades and, more particularly, to a turbine blade airfoil with serial impingement cooling through internal cavity-forming ribs.
  • Impingement cooling has been used in high pressure turbine stage nozzle vanes and rotor blades (hereinafter referred to as turbine blades) due to its high impingement heat transfer coefficient.
  • cooling air flows into and upwardly through the blade shank, through internal serpentine cavities and passages formed in the airfoil, and out through exit holes in the wall of the blade airfoil.
  • Impingement cooling was accomplished by inserting impingement inserts into the cavities of the airfoil.
  • Impingement inserts typically have hollow configurations generally conforming to the interior contour of the respective cavity of the airfoil but in spaced relation to the airfoil wall.
  • the inserts have jet air flow producing apertures in preselected locations.
  • High pressure cooling air from the turbine compressor is directed into the inserts through the blade shank in a well known manner and is exhausted through such apertures to form jets of air striking the interior surfaces of the airfoil wall for impingement cooling.
  • impingement inserts vibrate inside the blade and create metal fatigue.
  • the post impingement flows are usually permitted to bleed out of the airfoil right after the impingement and are used for film cooling.
  • this practice of using impingement inserts fails to recapture any post impingement air flow and use it for more cooling before it is allowed to bleed out of the airfoil.
  • the present invention provides a serial impingement cooling arrangement designed to satisfy the aforementioned need.
  • a turbine blade airfoil incorporates the serial impingement cooling arrangement in the internal cavity-forming ribs of the airfoil which both takes advantage of impingement air flow and recaptures the post impingement air flow and uses it for more cooling before it is allowed to bleed out of the airfoil.
  • the present invention is set forth in a turbine blade airfoil having leading and trailing edges and opposite side walls defining pressure and suction sides and merging together at said leading and trailing edges to define a hollow interior chamber for communication of cooling air flow to said side walls of said airfoil.
  • the airfoil also has a plurality of holes through the leading and trailing edges to permit exit of cooling air from the hollow interior chamber of the airfoil.
  • the present invention is directed to an impingement cooling arrangement which comprises: (a) a multiplicity of interior transverse walls spaced one to the next along the direction of a chord extending between the leading and trailing edges of the airfoil and disposed across the hollow interior chamber and rigidly connected with the opposite side walls so as to define a plurality of interior cavities in the chamber serially-arranged along the chord between the leading and trailing edges; and (b) means defining a pair of jet-producing orifices through each of the transverse walls for providing communication from one cavity to the next.
  • Pairs of orifices in a first plurality of the transverse walls that define a first series of cavities have respective pairs of axes which diverge from one another in a first direction of cooling air flow toward the leading edge of the airfoil.
  • the divergent relation of the orifice axes cause successive impingement against portions of the opposite side walls of successive cavities of the first series of cavities by portions of cooling air flow through the pairs of orifices in the first plurality of transverse walls before exiting from the airfoil through the exit holes in the leading edge of the airfoil.
  • Pairs of orifices in a second plurality of the transverse walls that define a second series of cavities have respective pairs of axes which diverge from one another in a second direction of cooling air flow, opposite from the first direction, toward the trailing edge of said airfoil.
  • the divergent relation of the orifice axes cause successive impingement against portions of the opposite side walls of successive cavities of the second series of cavities by portions of cooling air flow through the pairs of orifices in the second plurality of transverse walls before exiting from the airfoil through the exit holes in the trailing edge of the airfoil.
  • Fig. 1 is a perspective view of a prior art turbine engine blade having holes in the blade airfoil for exit of cooling air therefrom.
  • Fig. 2 is an enlarged cross-sectional view of the prior art blade airfoil taken along line 2--2 of Fig. 1.
  • Fig. 3 is an enlarged longitudinal sectional view of the prior art blade taken along line 3--3 of Fig. 1.
  • Fig. 4 is a view similar to Fig. 3 but now illustrating the serial impingement cooling arrangement of the present invention.
  • Fig. 5 is a diagrammatic view of of an illustration of a cooling air flow circuit defined by the serial impingement cooling arrangement of the present invention in the blade airfoil of Fig. 4.
  • Fig. 6 is an enlarged fragmentary view of the blade taken along line 6--6 of Fig. 4.
  • the hollow blade 10 includes an airfoil 12 having pressure and suction sides 14, 16 and leading and trailing edges 18, 20, and a base 22 mounting the airfoil 12 to a rotor (not shown) of the engine (not shown).
  • the base 22 has a platform 24 rigidly mounting the airfoil 12 and a dovetail root 26 for attaching the blade 10 to the rotor.
  • the airfoil 12 has opposite side walls 28, 30 defining the pressure and suction sides 14, 16 of the airfoil 12 which merge together at the leading and trailing edges 18, 20 of the airfoil 12 and are rigidly attached upright on the platform 24.
  • the airfoil 12 also has an end cap 32 which closes the outer ends of the side walls 28, 30.
  • the side walls 28, 30 and end cap 32 of the airfoil 12 contain small apertures or holes 34 which permit passage and exit of cooling air from the interior of the blade airfoil 12.
  • the airfoil 12 includes a plurality of interior spaced ribs or transverse walls 36 which extend across the hollow interior of the airfoil 12 and rigidly interconnected with the opposite side walls 28, 30 so as to define a series of interior cavities 38 in the airfoil 12 in a hollow interior chamber 40 of the airfoil.
  • the ribs 36 extend vertically and alternately connect to and terminate short of the end cap 32 at their upper ends and of a solid portion 42 of the base 22.
  • the ribs 36 define serpentine arrangements of cavities and passages within the interior of the airfoil 12 causing cooling air to flow along internal serpentine paths, as illustrated in Fig. 3, and exit through the holes 34 in the side walls 28, 30, leading and trailing edges 18, 20 and end cap 32 of the airfoil 12.
  • an arrangement of air flow jet-producing orifices 44 is provided in the transverse walls 46 of the turbine blade 48 for producing serial impingement cooling of the side walls 50, 52 of the blade 48. Otherwise the turbine blade 48 of Fig. 4 is the same as the turbine blade 10 of Figs. 1-3.
  • a pair of the orifices 44 are formed through each of the transverse walls 46 for providing communication from one interior cavity 38 to the next.
  • a first plurality of the transverse walls 46A, 46B that define a first series of cavities 38A, 38B have pairs of orifices 44 with respective pairs of axes 54A, 54B which diverge from one another in a first direction of cooling air flow from an intermediate one of the cavities 38C toward the leading edge 56 of the airfoil 58.
  • a second plurality of the transverse walls 46C, 46D, 46E that define a second series of cavities 38D, 38E, 38F have pairs of orifices 44 with respective pairs of axes 54C, 54D, 54E which diverge from one another in a second direction of cooling air flow, opposite from the first direction, from another intermediate one of the cavities 38G toward the trailing edge 62 of the airfoil 58.
  • the serial impingement airfoil 58 of Fig. 4 thus has two circuits, a forwardly directed one 66 and a rearwardly directed one 68, as seen in Fig. 5.
  • the number of circuit branches can be varied depending on the design.
  • the impingement orifices 44 are drilled on the cavity ribs or transverse walls 46 and oriented to directly impinge on either the pressure or the suction side wall surfaces.
  • the post impingement air will flow through the succeeding impingement holes to impinge on the surfaces of the next cavity 38 without creating the cross flow penalty in the same cavity.
  • This design will allow the further usage of post impingement air before it bleeds out of the airfoil 58 through the exit holes 60, 64.
  • the air can be bled out of the airfoil for either film cooling or recirculated for regenerative purposes.
  • the impingement orifices 44 can be either cast or drilled during the fabrication process.
  • the impingement transverse walls between cavities not only provide the impingement purpose but also continue to serve as the airfoil structure to carry the mechanical and thermal loads.
EP91308047A 1990-09-06 1991-09-03 Turbinenschaufel mit serieller Stossstrahlkühlung durch interne kammerformende Rippen Withdrawn EP0475658A1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US57816490A 1990-09-06 1990-09-06
US578164 1990-09-06

Publications (1)

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EP0475658A1 true EP0475658A1 (de) 1992-03-18

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Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0698724A3 (de) * 1994-08-23 1996-11-13 Gen Electric Kühlkreislauf des abströmendes einer Turbinenleitschaufel
WO1997006367A1 (de) * 1995-08-05 1997-02-20 Aloys Wobben Enteisen eines rotorblattes einer windkraftanlage
EP1001135A2 (de) * 1998-11-16 2000-05-17 General Electric Company Turbinenschaufel mit serieller Prallkühlung
DE19921644A1 (de) * 1999-05-10 2000-11-16 Abb Alstom Power Ch Ag Kühlbare Schaufel für eine Gasturbine
WO2004036038A1 (en) * 2002-10-17 2004-04-29 Lorenzo Battisti Anti-icing system for wind turbines
US7137784B2 (en) 2001-12-10 2006-11-21 Alstom Technology Ltd Thermally loaded component
EP1728970A2 (de) 2005-05-31 2006-12-06 United Technologies Corporation Kühlsystem für Turbinenschaufel
JP2011208625A (ja) * 2010-03-31 2011-10-20 Hitachi Ltd ガスタービン翼
WO2013085878A1 (en) * 2011-12-06 2013-06-13 Siemens Energy, Inc. Turbine blade incorporating trailing edge cooling design
KR101464988B1 (ko) * 2013-11-12 2014-11-26 연세대학교 산학협력단 냉각 성능 향상을 위한 내부유로 구조를 포함하는 가스터빈 블레이드
JP2015511678A (ja) * 2012-03-22 2015-04-20 アルストム テクノロジー リミテッドＡＬＳＴＯＭ Ｔｅｃｈｎｏｌｏｇｙ Ｌｔｄ タービン翼
CN107013252A (zh) * 2015-12-09 2017-08-04 通用电气公司 物件和冷却物件的方法
US10018052B2 (en) 2012-12-28 2018-07-10 United Technologies Corporation Gas turbine engine component having engineered vascular structure
US10036258B2 (en) 2012-12-28 2018-07-31 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10077664B2 (en) 2015-12-07 2018-09-18 United Technologies Corporation Gas turbine engine component having engineered vascular structure
US10094287B2 (en) 2015-02-10 2018-10-09 United Technologies Corporation Gas turbine engine component with vascular cooling scheme
US10221694B2 (en) 2016-02-17 2019-03-05 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US20190101009A1 (en) * 2017-10-03 2019-04-04 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626734B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10633980B2 (en) 2017-10-03 2020-04-28 United Technologies Coproration Airfoil having internal hybrid cooling cavities
US10704398B2 (en) 2017-10-03 2020-07-07 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities
US10774653B2 (en) 2018-12-11 2020-09-15 Raytheon Technologies Corporation Composite gas turbine engine component with lattice structure

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3111302A (en) * 1960-01-05 1963-11-19 Rolls Royce Blades for fluid flow machines
DE1946535A1 (de) * 1968-09-27 1970-04-23 Gen Electric Stroemungsfilmkuehlung fuer Bauteile von Gasturbinentriebwerken
FR2311176A1 (fr) * 1975-05-16 1976-12-10 Bbc Brown Boveri & Cie Ailette refroidie de turbine
GB2104965A (en) * 1981-08-31 1983-03-16 Gen Electric Multiple-impingement cooled structure
JPS59200001A (ja) * 1983-04-28 1984-11-13 Toshiba Corp ガスタ−ビン翼

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3111302A (en) * 1960-01-05 1963-11-19 Rolls Royce Blades for fluid flow machines
DE1946535A1 (de) * 1968-09-27 1970-04-23 Gen Electric Stroemungsfilmkuehlung fuer Bauteile von Gasturbinentriebwerken
FR2311176A1 (fr) * 1975-05-16 1976-12-10 Bbc Brown Boveri & Cie Ailette refroidie de turbine
GB2104965A (en) * 1981-08-31 1983-03-16 Gen Electric Multiple-impingement cooled structure
JPS59200001A (ja) * 1983-04-28 1984-11-13 Toshiba Corp ガスタ−ビン翼

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
PATENT ABSTRACTS OF JAPAN, Vol. 9, No. 67 (M-366)[1790] 27 March 1985; & JP-A-59 200 001 (TOSHIBA) 13 November 1984, Abstract. *

Cited By (37)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0698724A3 (de) * 1994-08-23 1996-11-13 Gen Electric Kühlkreislauf des abströmendes einer Turbinenleitschaufel
WO1997006367A1 (de) * 1995-08-05 1997-02-20 Aloys Wobben Enteisen eines rotorblattes einer windkraftanlage
EP1001135A2 (de) * 1998-11-16 2000-05-17 General Electric Company Turbinenschaufel mit serieller Prallkühlung
EP1001135A3 (de) * 1998-11-16 2001-12-05 General Electric Company Turbinenschaufel mit serieller Prallkühlung
DE19921644A1 (de) * 1999-05-10 2000-11-16 Abb Alstom Power Ch Ag Kühlbare Schaufel für eine Gasturbine
US6347923B1 (en) 1999-05-10 2002-02-19 Alstom (Switzerland) Ltd Coolable blade for a gas turbine
DE19921644B4 (de) * 1999-05-10 2012-01-05 Alstom Kühlbare Schaufel für eine Gasturbine
US7137784B2 (en) 2001-12-10 2006-11-21 Alstom Technology Ltd Thermally loaded component
US7637715B2 (en) 2002-10-17 2009-12-29 Lorenzo Battisti Anti-icing system for wind turbines
WO2004036038A1 (en) * 2002-10-17 2004-04-29 Lorenzo Battisti Anti-icing system for wind turbines
CN100359161C (zh) * 2002-10-17 2008-01-02 洛伦佐·巴蒂斯蒂 含有除冰和抗冰装备的风能转化系统以及防止和消除在风能转化系统的转子叶片上的冰层的方法
EP1728970A2 (de) 2005-05-31 2006-12-06 United Technologies Corporation Kühlsystem für Turbinenschaufel
EP1728970A3 (de) * 2005-05-31 2009-12-09 United Technologies Corporation Kühlsystem für Turbinenschaufel
JP2011208625A (ja) * 2010-03-31 2011-10-20 Hitachi Ltd ガスタービン翼
WO2013085878A1 (en) * 2011-12-06 2013-06-13 Siemens Energy, Inc. Turbine blade incorporating trailing edge cooling design
US9004866B2 (en) 2011-12-06 2015-04-14 Siemens Aktiengesellschaft Turbine blade incorporating trailing edge cooling design
CN104254669A (zh) * 2011-12-06 2014-12-31 西门子公司 包括后缘冷却设计的涡轮叶片
JP2015511678A (ja) * 2012-03-22 2015-04-20 アルストム テクノロジー リミテッドＡＬＳＴＯＭ Ｔｅｃｈｎｏｌｏｇｙ Ｌｔｄ タービン翼
US10662781B2 (en) 2012-12-28 2020-05-26 Raytheon Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10018052B2 (en) 2012-12-28 2018-07-10 United Technologies Corporation Gas turbine engine component having engineered vascular structure
US10036258B2 (en) 2012-12-28 2018-07-31 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10156359B2 (en) 2012-12-28 2018-12-18 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10570746B2 (en) 2012-12-28 2020-02-25 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10731473B2 (en) 2012-12-28 2020-08-04 Raytheon Technologies Corporation Gas turbine engine component having engineered vascular structure
KR101464988B1 (ko) * 2013-11-12 2014-11-26 연세대학교 산학협력단 냉각 성능 향상을 위한 내부유로 구조를 포함하는 가스터빈 블레이드
US10094287B2 (en) 2015-02-10 2018-10-09 United Technologies Corporation Gas turbine engine component with vascular cooling scheme
US10077664B2 (en) 2015-12-07 2018-09-18 United Technologies Corporation Gas turbine engine component having engineered vascular structure
CN107013252A (zh) * 2015-12-09 2017-08-04 通用电气公司 物件和冷却物件的方法
US10221694B2 (en) 2016-02-17 2019-03-05 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US20190101009A1 (en) * 2017-10-03 2019-04-04 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10633980B2 (en) 2017-10-03 2020-04-28 United Technologies Coproration Airfoil having internal hybrid cooling cavities
US10626734B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US10704398B2 (en) 2017-10-03 2020-07-07 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities
US10626733B2 (en) 2017-10-03 2020-04-21 United Technologies Corporation Airfoil having internal hybrid cooling cavities
US11649731B2 (en) 2017-10-03 2023-05-16 Raytheon Technologies Corporation Airfoil having internal hybrid cooling cavities
US10774653B2 (en) 2018-12-11 2020-09-15 Raytheon Technologies Corporation Composite gas turbine engine component with lattice structure
US11168568B2 (en) 2018-12-11 2021-11-09 Raytheon Technologies Corporation Composite gas turbine engine component with lattice

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