EP0032646A1 - Gasturbinenleitschaufel - Google Patents

Gasturbinenleitschaufel Download PDF

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Publication number
EP0032646A1
EP0032646A1 EP80401849A EP80401849A EP0032646A1 EP 0032646 A1 EP0032646 A1 EP 0032646A1 EP 80401849 A EP80401849 A EP 80401849A EP 80401849 A EP80401849 A EP 80401849A EP 0032646 A1 EP0032646 A1 EP 0032646A1
Authority
EP
European Patent Office
Prior art keywords
jacket
turbine distributor
perforations
rows
leading edge
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
EP80401849A
Other languages
English (en)
French (fr)
Other versions
EP0032646B1 (de
Inventor
Denis René Guy Laffitte
Guy Henri Louis
Alain Marius Marcel Raybaud
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.)
Safran Aircraft Engines SAS
Original Assignee
Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
SNECMA SAS
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 Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA, SNECMA SAS filed Critical Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Publication of EP0032646A1 publication Critical patent/EP0032646A1/de
Application granted granted Critical
Publication of EP0032646B1 publication Critical patent/EP0032646B1/de
Expired 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall

Definitions

  • the present invention relates to a turbine distributor blade.
  • High performance turbomachines are equipped with turbine guide vanes capable of withstanding temperatures in the region of 1500 ° C. and it is even envisaged to use blades capable of operating at higher temperatures.
  • Such blades require a very sophisticated cooling system and a system of internal channels. It is known to use turbine distributor vanes comprising at least one internal cavity in which a jacket of perforated sheet metal bears against the walls by means of projecting elements.
  • the arrangement of the projecting elements constituted by cylindrical studs and fins does not generally make it possible to obtain a coefficient of heat exchange sufficient for the operating temperatures envisaged.
  • the projecting elements against which the jacket is supported are arranged for the lower and upper surfaces in an area extending substantially over the entire height to become minimal at the top, the internal cavity having an opening at its upper part through which the coolant enters.
  • This arrangement of the projecting elements according to the invention which consists of arranging them in the form of triangle or trapezoidal zones, makes it possible to arrange evolving passage sections such as the ratio of the residual cold air flow rate over the passage area remains substantially constant.
  • the jacket is kept under pressure against the projecting elements and in the open position by a rigid locking device such as a longitudinal lock pierced with orifices.
  • a rigid locking device such as a longitudinal lock pierced with orifices.
  • the main chamber located behind the leading edge and in which the jacket is mounted is divided into three cooling zones which do not communicate with each other, which improves the cooling of the blade by the air circulation.
  • FIG. 1 shows a part of the turbine of a turbomachine which is located at the outlet of a combustion chamber 1 and which comprises a guide vane 2 and a turbine fin 3 which are located in the channel d annular flow 4 of the combustion gases.
  • the guide blade 2 is part of a row of blades which is arranged circularly in the flow channel 4.
  • Each blade 2 has in known manner a head 5 and a foot 6 ( Figures 1, 2, 3) said head having an opening 7 through which penetrates the cooling air from the compressor. The air is distributed in the various internal channels as will be described later.
  • Each blade comprises a main chamber 6, an intermediate chamber 9 and a trailing edge area 10, also visible in FIGS. 4 and 5.
  • the main chamber 8 occupies almost 2/3 of the internal volume of the blade; this makes it possible to reduce the Mach number of the heat transfer fluid, therefore to maintain a high pressure level.
  • the latter is first channeled in an internal jacket 12 made of sheet metal which insulates it from the walls.
  • the sheet metal jacket 12 consists of two plates, one of which 12a extends over the entire width of the main chamber 8 and the other of which 12b is fixed on the first to form a U-shaped jacket open to one of its ends towards the leading edge 11.
  • the jacket 12 is supported in substantially sealed manner by the plate 12a against the central partition 13 separating the main chamber 8 from the intermediate chamber 9 and it is supported by its plate 12a against cylindrical or frustoconical studs 14 located on the upper surface side and by its plate 12b against transverse fins 15 located on the lower side.
  • the plates 12a and 12b of the jacket rest on two ribs 16, 16a and they are held in the open position by a longitudinal latch 17 engaged in ribs provided in the edges of the two plates 12a, 12b .
  • the latch 17 consists of a plate having openings 17a opening towards the leading edge 11.
  • the rows of studs 14 coming from the foundry on the internal face of the upper surface as well as the transverse fins 15 formed on the lower surface are arranged in an area extending substantially over the entire height. and whose width which is maximum at the base gradually evolves over the entire height to become minimum at the top.
  • This arrangement of the projecting elements 14 and 15 which consists of arranging them in the shape of a triangle or trapezoid makes it possible to arrange scalable passage sections such that the ratio of the flow of residual cold air over the passage area remains substantially constant . This results in better cooling of the walls of the blade.
  • the blade has rows of perforators 18 on the leading edge, of perforations 19 on the lower surface, near the leading edge and near the internal partition, of perforations 20 on the upper surface near the leading edge.
  • the diameter of these perforations is very small, of the order of 0.3 mm.
  • a staggered arrangement will be adopted.
  • the arrangement of the jacket 12 in the main bedroom is such that it divides said bedroom into three independent zones A, B, C (see FIG. 5).
  • zone B The air which arrives through the head 7 of the blade 2 is also distributed on the one hand in zone B, passes through the fins 15 and escapes through the perforations 19, and on the other hand in zone C, passes through the studs 14 and escapes through the perforations 20.
  • zone B between the sleeve 12 and the fins 15 and in zone C between the sleeve 12 and the studs 14 is small; this allows a high Mach number and a low feed rate which is favorable for convection cooling.
  • the external exchange coefficient (intake of calories) being stronger on the upper surface than on the lower surface, we use fins on the lower surface and studs on the upper surface.
  • This cooling by fins is less effective but provides a lower pressure drop.
  • the pads 14 have a larger wetted surface and create a turbulence favorable to exchanges.
  • the zones B for supplying the fins 15 and C for supplying the studs 14 decrease as one approaches the foot of the dawn. Conversely, the length of the fins or rows of studs increases, which makes it possible to largely balance the exchanges over the entire surface of the blade.
  • the intermediate chamber 9 comprises on the lower surface side (FIG. 2) a smooth part 21 which occupies a right triangle whose base is constituted by the upper part of the chamber and the top by the internal lower corner.
  • a smooth part 21 which occupies a right triangle whose base is constituted by the upper part of the chamber and the top by the internal lower corner.
  • rows of studs 22 from the foundry are provided, distributed over a right triangle whose apex occupies the upper downstream corner of the chamber 9.
  • the longitudinal ribs 23 and the rows of studs 24 are sufficient for cooling the upper surface. It should also be noted that, in this zone, the density of the fins decreases and that the density of the studs increases when one approaches the foot of the dawn.
  • the intermediate chamber 9 opens into a groove 26 (FIG. 4, 5) occupying the entire length of the trailing edge, by slots 27 separated by bridges 28 from the foundry.
  • the slots 27 are divided into passages 29, 30, 31 delimited by two rows of studs 32, 33 secured to the lower surface as well as the upper surface.
  • a latch 17 has been shown consisting of a perforated plate to hold the edges of the jacket 12 in the open position, it is also possible to keep the plates open by engaging the ends of the plates 12a and 12b in slots formed in the ribs 16, 16a.
EP80401849A 1980-01-10 1980-12-23 Gasturbinenleitschaufel Expired EP0032646B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8000458A FR2473621A1 (fr) 1980-01-10 1980-01-10 Aube de distributeur de turbine
FR8000458 1980-01-10

Publications (2)

Publication Number Publication Date
EP0032646A1 true EP0032646A1 (de) 1981-07-29
EP0032646B1 EP0032646B1 (de) 1984-06-13

Family

ID=9237402

Family Applications (1)

Application Number Title Priority Date Filing Date
EP80401849A Expired EP0032646B1 (de) 1980-01-10 1980-12-23 Gasturbinenleitschaufel

Country Status (5)

Country Link
US (1) US4403917A (de)
EP (1) EP0032646B1 (de)
JP (1) JPS56138403A (de)
DE (1) DE3068276D1 (de)
FR (1) FR2473621A1 (de)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2119028A (en) * 1982-04-27 1983-11-09 Rolls Royce Aerofoil for a gas turbine engine
EP0140257A1 (de) * 1983-10-28 1985-05-08 Westinghouse Electric Corporation Anordnung für die Kühlung der Hinterkante einer Leitschaufel
FR2653171A1 (fr) * 1989-10-18 1991-04-19 Snecma Carter de compresseur de turbomachine muni d'un dispositif de pilotage de son diametre interne.
GB2261032A (en) * 1991-08-23 1993-05-05 Mitsubishi Heavy Ind Ltd Gas turbine blade with skin and core construction
EP1849960A2 (de) * 2006-04-27 2007-10-31 Hitachi, Ltd. Turbinenschaufel mit innerem Kühlkanal
EP3023586A1 (de) * 2014-11-21 2016-05-25 Siemens Aktiengesellschaft Hohlschaufelkörper, Einsteckrippe und Hohlschaufel

Families Citing this family (39)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2725474B1 (fr) * 1984-03-14 1996-12-13 Snecma Aube de distributeur de turbine refroidie
GB2170867B (en) * 1985-02-12 1988-12-07 Rolls Royce Improvements in or relating to gas turbine engines
DE3685852T2 (de) * 1985-04-24 1992-12-17 Pratt & Whitney Canada Turbinenmotor mit induziertem vordrall am kompressoreinlass.
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
JP2862536B2 (ja) * 1987-09-25 1999-03-03 株式会社東芝 ガスタービンの翼
US5281084A (en) * 1990-07-13 1994-01-25 General Electric Company Curved film cooling holes for gas turbine engine vanes
JPH0559718U (ja) * 1992-01-21 1993-08-06 アサヒ通信株式会社 ケーブルの構造
US5328331A (en) * 1993-06-28 1994-07-12 General Electric Company Turbine airfoil with double shell outer wall
US5484258A (en) * 1994-03-01 1996-01-16 General Electric Company Turbine airfoil with convectively cooled double shell outer wall
US5516260A (en) * 1994-10-07 1996-05-14 General Electric Company Bonded turbine airfuel with floating wall cooling insert
FR2743391B1 (fr) 1996-01-04 1998-02-06 Snecma Aube refrigeree de distributeur de turbine
US5601399A (en) * 1996-05-08 1997-02-11 Alliedsignal Inc. Internally cooled gas turbine vane
US5772397A (en) * 1996-05-08 1998-06-30 Alliedsignal Inc. Gas turbine airfoil with aft internal cooling
GB2350867B (en) * 1999-06-09 2003-03-19 Rolls Royce Plc Gas turbine airfoil internal air system
US6742987B2 (en) 2002-07-16 2004-06-01 General Electric Company Cradle mounted turbine nozzle
EP1589192A1 (de) * 2004-04-20 2005-10-26 Siemens Aktiengesellschaft Turbinenschaufel mit einem Prallkühleinsatz
US20080145208A1 (en) * 2006-12-19 2008-06-19 General Electric Company Bullnose seal turbine stage
US7578653B2 (en) 2006-12-19 2009-08-25 General Electric Company Ovate band turbine stage
JP2009162119A (ja) * 2008-01-08 2009-07-23 Ihi Corp タービン翼の冷却構造
US8961133B2 (en) * 2010-12-28 2015-02-24 Rolls-Royce North American Technologies, Inc. Gas turbine engine and cooled airfoil
US9759072B2 (en) * 2012-08-30 2017-09-12 United Technologies Corporation Gas turbine engine airfoil cooling circuit arrangement
US9169733B2 (en) * 2013-03-20 2015-10-27 General Electric Company Turbine airfoil assembly
US20160222796A1 (en) * 2013-09-18 2016-08-04 United Technologies Corporation Manufacturing method for a baffle-containing blade
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
EP3189213A1 (de) 2014-09-04 2017-07-12 Siemens Aktiengesellschaft Internes kühlsystem mit einlageformenden wandnahen kühlkanälen in einem hinteren kühlhohlraum einer gasturbinenschaufel
EP3167160A1 (de) * 2014-09-04 2017-05-17 Siemens Aktiengesellschaft Internes kühlsystem mit einsatz zur bildung von wandnahen kühlkanälen in einem hinteren kühlhohlraum einer gasturbinenschaufel mit wärmeableitenden rippen
US9840930B2 (en) 2014-09-04 2017-12-12 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil
US9879554B2 (en) * 2015-01-09 2018-01-30 Solar Turbines Incorporated Crimped insert for improved turbine vane internal cooling
WO2016148693A1 (en) 2015-03-17 2016-09-22 Siemens Energy, Inc. Internal cooling system with converging-diverging exit slots in trailing edge cooling channel for an airfoil in a turbine engine
US10436048B2 (en) * 2016-08-12 2019-10-08 General Electric Comapny Systems for removing heat from turbine components
US10364685B2 (en) * 2016-08-12 2019-07-30 Gneral Electric Company Impingement system for an airfoil
US10443397B2 (en) * 2016-08-12 2019-10-15 General Electric Company Impingement system for an airfoil
US10408062B2 (en) * 2016-08-12 2019-09-10 General Electric Company Impingement system for an airfoil
CN109404051B (zh) * 2018-12-29 2021-10-26 中国科学院工程热物理研究所 一种涡轮导向器的浮动定位及传扭结构
US11085374B2 (en) * 2019-12-03 2021-08-10 General Electric Company Impingement insert with spring element for hot gas path component
US11428166B2 (en) * 2020-11-12 2022-08-30 Solar Turbines Incorporated Fin for internal cooling of vane wall
US11898463B2 (en) * 2021-03-29 2024-02-13 Rtx Corporation Airfoil assembly with fiber-reinforced composite rings
US11549378B1 (en) 2022-06-03 2023-01-10 Raytheon Technologies Corporation Airfoil assembly with composite rings and sealing shelf
US20230417146A1 (en) * 2022-06-23 2023-12-28 Solar Turbines Incorporated Pneumatically variable turbine nozzle

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2150476A1 (de) * 1971-08-25 1973-04-06 Rolls Royce

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US3017159A (en) * 1956-11-23 1962-01-16 Curtiss Wright Corp Hollow blade construction
FR1374159A (fr) * 1962-12-05 1964-10-02 Gen Motors Corp Pale de turbine
US3370829A (en) * 1965-12-20 1968-02-27 Avco Corp Gas turbine blade construction
US3475107A (en) * 1966-12-01 1969-10-28 Gen Electric Cooled turbine nozzle for high temperature turbine
BE755567A (fr) * 1969-12-01 1971-02-15 Gen Electric Structure d'aube fixe, pour moteur a turbines a gaz et arrangement de reglage de temperature associe
US3647316A (en) * 1970-04-28 1972-03-07 Curtiss Wright Corp Variable permeability and oxidation-resistant airfoil
US3635587A (en) * 1970-06-02 1972-01-18 Gen Motors Corp Blade cooling liner
GB1304678A (de) * 1971-06-30 1973-01-24
SU364747A1 (ru) * 1971-07-08 1972-12-28 Охлаждаемая лопатка турбол1ашины
US4025226A (en) * 1975-10-03 1977-05-24 United Technologies Corporation Air cooled turbine vane
US4063851A (en) * 1975-12-22 1977-12-20 United Technologies Corporation Coolable turbine airfoil
GB1565361A (en) * 1976-01-29 1980-04-16 Rolls Royce Blade or vane for a gas turbine engien
GB2017229B (en) * 1978-03-22 1982-07-14 Rolls Royce Guides vanes for gas turbine enginess

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2150476A1 (de) * 1971-08-25 1973-04-06 Rolls Royce

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2119028A (en) * 1982-04-27 1983-11-09 Rolls Royce Aerofoil for a gas turbine engine
EP0140257A1 (de) * 1983-10-28 1985-05-08 Westinghouse Electric Corporation Anordnung für die Kühlung der Hinterkante einer Leitschaufel
FR2653171A1 (fr) * 1989-10-18 1991-04-19 Snecma Carter de compresseur de turbomachine muni d'un dispositif de pilotage de son diametre interne.
EP0424253A1 (de) * 1989-10-18 1991-04-24 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Turbokompressorgehäuse mit Spielkontrollvorrichtung für den inneren Durchmesser
GB2261032A (en) * 1991-08-23 1993-05-05 Mitsubishi Heavy Ind Ltd Gas turbine blade with skin and core construction
GB2261032B (en) * 1991-08-23 1995-04-05 Mitsubishi Heavy Ind Ltd Rotor blade for a gas turbine
EP1849960A2 (de) * 2006-04-27 2007-10-31 Hitachi, Ltd. Turbinenschaufel mit innerem Kühlkanal
EP1849960A3 (de) * 2006-04-27 2010-03-10 Hitachi, Ltd. Turbinenschaufel mit innerem Kühlkanal
EP3023586A1 (de) * 2014-11-21 2016-05-25 Siemens Aktiengesellschaft Hohlschaufelkörper, Einsteckrippe und Hohlschaufel
WO2016078851A1 (de) * 2014-11-21 2016-05-26 Siemens Aktiengesellschaft Hohlschaufelkörper, einsteckrippe und hohlschaufel

Also Published As

Publication number Publication date
JPS56138403A (en) 1981-10-29
EP0032646B1 (de) 1984-06-13
FR2473621A1 (fr) 1981-07-17
JPS6148609B2 (de) 1986-10-24
DE3068276D1 (en) 1984-07-19
FR2473621B1 (de) 1983-05-13
US4403917A (en) 1983-09-13

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