US20040109763A1 - Method of producing and assembling a cooling device inside an axial-flow gas turbine blade, and axial-flow gas turbine blade produced using such a method - Google Patents

Method of producing and assembling a cooling device inside an axial-flow gas turbine blade, and axial-flow gas turbine blade produced using such a method Download PDF

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Publication number
US20040109763A1
US20040109763A1 US10/604,337 US60433703A US2004109763A1 US 20040109763 A1 US20040109763 A1 US 20040109763A1 US 60433703 A US60433703 A US 60433703A US 2004109763 A1 US2004109763 A1 US 2004109763A1
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US
United States
Prior art keywords
chamber
blade
openings
insert
forcing
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.)
Abandoned
Application number
US10/604,337
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English (en)
Inventor
Paolo Ciacci
Daniele Coutandin
Domenico Dalle Crode
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.)
GE Avio SRL
Original Assignee
Avio SpA
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 Avio SpA filed Critical Avio SpA
Publication of US20040109763A1 publication Critical patent/US20040109763A1/en
Abandoned 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
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • 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
    • 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a method of producing and assembling a cooling device inside an axial-flow gas turbine blade, and in particular an adjustable-angle blade of a variable-geometry gas turbine stator, to which the following description refers purely by way of example.
  • the airfoil profile of such a blade is hinged to the annular platforms defining the gas conduit in the stator, and comprises a tail portion connected in sliding manner to the platforms.
  • a method of producing and assembling a cooling device inside a blade of an axial-flow gas turbine comprising an airfoil profile having an inner surface defining a chamber, and two connecting end portions located on opposite sides of said airfoil profile for connection to respective supporting structures forming part of said turbine, and having respective openings for the passage of a cooling fluid and which come out inside said chamber; the method comprising the steps of forming an insert having a number of holes; and positioning said insert inside said chamber so as to face said inner surface and direct a relative stream of said cooling fluid through each said hole on to said inner surface; characterized in that said insert is formed by producing a first and at least a second body separate from each other and each of a size approximating but no larger than that of at least one of said openings; and in that positioning said insert inside said chamber comprises the step of inserting said first and said second body successively through said openings.
  • the present invention also relates to an axial-flow gas turbine blade.
  • a blade for an axial-flow gas turbine comprising an airfoil profile having an inner surface defining a chamber; two connecting end portions located on opposite sides of said airfoil profile for connection to respective structures forming part of said turbine, and having respective openings for the passage of a cooling fluid and which come out inside said chamber; and a cooling device comprising an insert having a number of holes and positioned inside said chamber so as to face said inner surface and direct a relative stream of said cooling fluid through each said hole on to said inner surface; characterized in that said insert comprises a first and at least a second body separate from each other and each of a size approximating but no larger than that of at least one of said openings, so as to be insertable through the openings.
  • FIG. 1 shows a schematic exploded side view of a preferred embodiment of the method according to the present invention for producing and assembling a cooling device inside a gas turbine blade;
  • FIGS. 2 and 3 show a cross section and a cutaway view in perspective respectively of a blade produced in accordance with the FIG. 1 method.
  • Number 1 in the accompanying drawings indicates a blade for a stator (not shown) of an aircraft variable-geometry axial-flow gas turbine (not shown).
  • Blade 1 comprises an airfoil profile 2 housed, in use, inside an annular gas conduit of the turbine, and in turn comprising a front portion 4 hinged about an axis 5 and by respective opposite pins 6 a, 6 b to two annular platforms (not shown) of the stator defining the conduit.
  • Pins 6 a, 6 b are coaxial, are formed in one piece with portion 4 , and define respective circular openings 7 a, 7 b, which are coaxial along axis 5 , come out inside a chamber 8 defined by an inner surface 9 of profile 2 , are relatively small in diameter with respect to the size of chamber 8 , and, in use, permit the passage of a stream of cooling air.
  • Profile 2 also comprises a tail portion 11 , in turn comprising a high-pressure wall 12 , a low-pressure wall 13 , and two walls 14 located on opposite axial sides of walls 12 , 13 , and connected in sliding manner, in use, to said platforms (FIG. 3).
  • Blade 1 also comprises a cooling device 15 , in turn comprising an assembly 16 of four separate bodies 17 , 18 , 19 , 20 housed inside chamber 8 .
  • Each body 17 , 18 , 19 , 20 comprises a relative portion of a lateral wall 21 (FIG. 3), which faces surface 9 , extends along the whole of surface 9 , and has a number of holes 22 (shown schematically) through which respective cooling air jets flow from the centre of chamber 8 on to surface 9 .
  • Bodies 17 , 18 , 19 , 20 are elongated parallel to axis 5 , are aligned with one another in a direction A radial with respect to axis 5 , and are each of a size, crosswise to axis 5 , approximating but no larger than the diameter of at least one of openings 7 a, 7 b, so as to be insertable axially through openings 7 a, 7 b.
  • Bodies 17 , 18 , 20 are box- or shell-shaped, and rest on relative inner ribs 23 (shown partly) of profile 2 , so as to be kept detached from surface 9 .
  • body 17 is located inside chamber 8 at the end of portion 11 , is substantially wedge-shaped, defines an inner cavity 24 , is interposed between walls 14 (FIG. 3), and is forced between walls 12 , 13 in direction A towards the trailing edge of portion 11 .
  • Body 18 is interposed between bodies 17 and 19 , rests axially on walls 14 , rests on body 17 in direction A, and defines an inner cavity 25 communicating with cavity 24 through two openings 26 a, 26 b formed in front of each other in respective bodies 17 , 18 (FIG. 2).
  • Body 20 is located inside chamber 8 close to the leading edge of portion 4 , has a substantially half-moon-shaped cross section, defines an inner cavity 28 , and rests axially on an inner shoulder 29 of pin 6 a (FIG. 2).
  • body 19 is tubular, and is bounded by a truncated-cone-shaped outer surface 31 resting, in direction A, on two concave surfaces 32 , 33 complementary to surface 31 and bounding respective bodies 18 , 20 .
  • Body 19 defines an inner channel 34 connecting openings 7 a, 7 b, and in turn communicating with cavity 25 through two openings 35 a, 35 b formed in front of each other in respective bodies 18 , 19 , and with cavity 28 through two openings 37 a, 37 b formed in front of each other in respective bodies 20 , 19 .
  • Body 19 comprises two opposite end portions 40 , 41 .
  • Portion 40 is housed in opening 7 a, rests on the inner surface 42 of pin 6 a, and is connected integrally to surface 42 by a brazed joint 42 a not shown in detail.
  • Portion 41 is connected to pin 6 b with the interposition of an annular retaining member 43 , which forms part of device 15 , is housed in opening 7 b, and comprises a cylindrical portion 44 connected integrally, preferably brazed, to pin 6 b in a manner not shown in detail.
  • Member 43 also comprises a tab 46 , which projects from portion 44 , perpendicularly to axis 5 , rests axially on portion 41 , and is connected integrally to portion 41 by a brazed joint 46 a not shown in detail.
  • device 15 also comprises two C-section spacers 48 , 49 , which are interposed axially between tab 46 and respective bodies 18 , 20 , and are deformed elastically to force bodies 18 , 20 elastically and axially towards pin 6 a.
  • Assembly 16 defines an insert or plate, which is externally substantially a negative of the shape of chamber 8 , and can be dismantled, i.e. into bodies 17 , 18 , 19 , 20 smaller than, and therefore insertable successively through, openings 7 a, 7 b.
  • body 17 is first inserted through opening 7 b and pushed to the end of the chamber towards the trailing edge of portion 11 .
  • Body 18 is then inserted through opening 7 b into chamber 8 and positioned adjacent to body 17 in direction A; body 20 is inserted through opening 7 a to rest on portion 4 in direction A; and finally, body 19 is inserted, and, as it moves along axis 5 , forces bodies 17 , 18 , 20 in direction A by virtue of the taper of surface 31 .
  • portion 40 is brazed to pin 6 a, and portion 41 is fixed to pin 6 b by attaching member 43 and interposing spacers 48 , 49 between member 43 and bodies 18 , 20 .
  • the method described therefore provides for inserting a cooling insert or plate easily inside profile 2 , even when the openings 7 a, 7 b in the connecting end portions of blade 1 are relatively small, by the insert or plate being dismantled into a number of separate parts (four in the example described).
  • Device 15 is also relatively easy to assemble and fix to profile 2 , by only body 19 being connected integrally to pins 6 a, 6 b, and by bodies 18 , 19 , 20 being locked automatically inside chamber 8 by body 19 and member 43 .
  • dismantling the insert or plate into at least two separate successively inserted bodies also applies advantageously to other types of blades having relatively small access openings with respect to the transverse dimensions of the inner chamber of the airfoil profile.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US10/604,337 2002-07-12 2003-07-11 Method of producing and assembling a cooling device inside an axial-flow gas turbine blade, and axial-flow gas turbine blade produced using such a method Abandoned US20040109763A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
IT2002TO000607A ITTO20020607A1 (it) 2002-07-12 2002-07-12 Metodo per la realizzazione ed il montaggio di un dispositivo di raffreddamento in una paletta di una turbina assiale a gas e paletta per un
ITTO2002A000607 2002-07-12

Publications (1)

Publication Number Publication Date
US20040109763A1 true US20040109763A1 (en) 2004-06-10

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US10/604,337 Abandoned US20040109763A1 (en) 2002-07-12 2003-07-11 Method of producing and assembling a cooling device inside an axial-flow gas turbine blade, and axial-flow gas turbine blade produced using such a method

Country Status (4)

Country Link
US (1) US20040109763A1 (it)
EP (1) EP1380725A3 (it)
CA (1) CA2435070A1 (it)
IT (1) ITTO20020607A1 (it)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20040096321A1 (en) * 2002-08-06 2004-05-20 Avio S.P.A. Variable-geometry turbine stator blade, particularly for aircraft engines
CN103261584A (zh) * 2010-12-22 2013-08-21 西门子公司 燃气涡轮机叶片或叶瓣的冲击冷却
US10422244B2 (en) 2015-03-16 2019-09-24 General Electric Company System for cooling a turbine shroud
US10612397B2 (en) * 2016-02-22 2020-04-07 Mitsubishi Hitachi Power Systems, Ltd. Insert assembly, airfoil, gas turbine, and airfoil manufacturing method

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2540969A1 (en) 2011-06-27 2013-01-02 Siemens Aktiengesellschaft Impingement cooling of turbine blades or vanes
EP2573325A1 (en) 2011-09-23 2013-03-27 Siemens Aktiengesellschaft Impingement cooling of turbine blades or vanes
EP2860348A1 (de) * 2013-10-08 2015-04-15 Siemens Aktiengesellschaft Einsatz für eine Turbinenschaufel aus mehreren Bauteilen und zugehöriges Verfahren
US10502070B2 (en) 2016-11-17 2019-12-10 United Technologies Corporation Airfoil with laterally insertable baffle

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4086021A (en) * 1976-01-19 1978-04-25 Stal-Laval Turbin Ab Cooled guide vane
US4257734A (en) * 1978-03-22 1981-03-24 Rolls-Royce Limited Guide vanes for gas turbine engines

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1605194A (en) * 1974-10-17 1983-04-07 Rolls Royce Rotor blade for gas turbine engines
US4798515A (en) * 1986-05-19 1989-01-17 The United States Of America As Represented By The Secretary Of The Air Force Variable nozzle area turbine vane cooling
US6543993B2 (en) * 2000-12-28 2003-04-08 General Electric Company Apparatus and methods for localized cooling of gas turbine nozzle walls

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4086021A (en) * 1976-01-19 1978-04-25 Stal-Laval Turbin Ab Cooled guide vane
US4257734A (en) * 1978-03-22 1981-03-24 Rolls-Royce Limited Guide vanes for gas turbine engines

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20040096321A1 (en) * 2002-08-06 2004-05-20 Avio S.P.A. Variable-geometry turbine stator blade, particularly for aircraft engines
US6913440B2 (en) * 2002-08-06 2005-07-05 Avio S.P.A. Variable-geometry turbine stator blade, particularly for aircraft engines
CN103261584A (zh) * 2010-12-22 2013-08-21 西门子公司 燃气涡轮机叶片或叶瓣的冲击冷却
US20130272896A1 (en) * 2010-12-22 2013-10-17 Anthony Davis Impingement cooling of gas turbine blades or vanes
CN103261584B (zh) * 2010-12-22 2015-06-17 西门子公司 涡轮机部件、置于其中空翼型内的冲击管及其组装方法
US9500087B2 (en) * 2010-12-22 2016-11-22 Siemens Aktiengesellschaft Impingement cooling of gas turbine blades or vanes
US10422244B2 (en) 2015-03-16 2019-09-24 General Electric Company System for cooling a turbine shroud
US10612397B2 (en) * 2016-02-22 2020-04-07 Mitsubishi Hitachi Power Systems, Ltd. Insert assembly, airfoil, gas turbine, and airfoil manufacturing method

Also Published As

Publication number Publication date
EP1380725A3 (en) 2004-09-15
CA2435070A1 (en) 2004-01-12
ITTO20020607A1 (it) 2004-01-12
EP1380725A2 (en) 2004-01-14
ITTO20020607A0 (it) 2002-07-12

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