EP0034961B1 - Perfectionnement aux aubes de turbines refroidies - Google Patents

Perfectionnement aux aubes de turbines refroidies Download PDF

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Publication number
EP0034961B1
EP0034961B1 EP19810400179 EP81400179A EP0034961B1 EP 0034961 B1 EP0034961 B1 EP 0034961B1 EP 19810400179 EP19810400179 EP 19810400179 EP 81400179 A EP81400179 A EP 81400179A EP 0034961 B1 EP0034961 B1 EP 0034961B1
Authority
EP
European Patent Office
Prior art keywords
blade
cavity
trailing edge
flow
orifices
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.)
Expired
Application number
EP19810400179
Other languages
German (de)
English (en)
French (fr)
Other versions
EP0034961A1 (fr
Inventor
Michel Léonard Cuvillier
Thierry Marie Charles Schindler
Jacques Philippe Henri Tirole
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
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 filed Critical Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Publication of EP0034961A1 publication Critical patent/EP0034961A1/fr
Application granted granted Critical
Publication of EP0034961B1 publication Critical patent/EP0034961B1/fr
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
    • 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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/126—Baffles or ribs
    • 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
    • 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
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the present invention relates to an improvement to the blades of cooled turbines.
  • document US-A-3017 159 shows how one can achieve, in a blade, a combination of radial and tangential circulations, but this arrangement does not allow good homogeneity in cooling because the fluid which circulates along the part internal of the leading edge is used, when it is already heated, to cool the external part of the trailing edge.
  • Document US-A-4 180 373 shows specialized compartments for the leading edge and the trailing edge but does not teach a method of homogeneous cooling of the latter by zone by zone control.
  • the envisaged arrangement does not allow the respective flow rates of the compartments to be adjusted.
  • the improved blade according to the invention which makes it possible to meet these criteria, comprises two cavities separated over the entire height of the blade by an internal partition supplied with a flow of cooling fluid including an upstream cavity and a divided downstream cavity in two radial and "tangential" flow zones towards the trailing edge.
  • the upstream cavity has fins favoring heat exchanges and holes drilled in the leading edge and the downstream cavity comprises, in the central zone for the radial flow of the fluid, bridges, that is to say bars connecting the two internal faces of large section, while in the trailing edge area with "tangential" flow, said downstream cavity also includes bridges but of smaller section supplemented by deflector members guiding the radial flow, tangentially toward the trailing edge holes.
  • This turbine blade according to the invention has a simple and efficient cooling circuit and is remarkable for its ease of manufacture and for the distribution of the heat exchanger members.
  • This blade is of a very easy realization, because it allows the use of a core with two cavities.
  • this blade has a high permeability, as well as the absence of baffles. Furthermore, by modifying the entry conditions (section of the diaphragms in particular), the exit conditions (apertures at the top, section and location of the emission zones) and by varying the size and spacing of the bridges, it is possible to control the local cooling of each portion of the blade, so avoid temperature heterogeneities in the metal of the blade.
  • the good general cooling conditions and this absence of hot spots mean that this blade can be adapted to temperature conditions of the hot gases which can be much more severe than those of the blades of the prior art.
  • FIGs 1 and 2 there is shown an embodiment of a turbine blade according to the invention which is cast in refractory metal and which comprises an upstream cavity 1 with radial flow and a downstream cavity 2 separated into two zones, one at radial flow and a tangential flow towards the trailing edge 3.
  • the upstream cavity 1 is delimited by a leading edge 4 and by a partition 5 and it is crossed by an air flow which moves according to the arrows F, by entering through an opening 6 made in the foot 7 of the blade which has a platform 23, said flow being evacuated in part by holes 8 drilled in the leading edge and in part by an orifice 9 formed in the dawn hat 10.
  • the hat 10 of the dawn is brought back or came from foundry with the dawn.
  • the holes 8 of the leading edge are drilled either on the upper surface part at 8a, or on the lower surface part at 8b,. and either at the end of the leading edge in 8c if the available cooling air pressure is sufficient.
  • the internal face of the leading edge 4 over the upper 2/3 of its height is provided with fins 11 promoting heat exchange fiques made up of small parallel ribs, extending in an axial plane. These ribs 11 may have the same height or different heights. It may also be advantageous to vary their spacing along the cavity.
  • the downstream cavity 2, delimited by the internal partition 5 and by the trailing edge 3, is divided into two zones including a central zone comprising bridges 12 of large section and a trailing edge zone comprising bridges 13 of small section arranged in staggered rows and deflector members 14, these divert part of the radia flow! from the cavity 2 towards the orifices 15 of the trailing edge 3. This part of the flow is thus deflected to pass from the substantially radial direction to the direction according to the arrows F2.
  • the air flow which enters the cavity 2 through the orifice 16 located at the foot of the blade, along the arrow F1 is divided into a radial flow moving in the central area comprising the bridges 12 and in a tangential flow moving in the area of the trailing edge comprising the bridges 13 of smaller section.
  • the radial air flow is evacuated through the orifice 17 provided in the cap 10 of the blade, while the tangential air flow is evacuated through the orifices 15 of the trailing edge which open either on the lower surface of the edge of leakage, either on the edge of the latter as shown in Figure 2.
  • the size and spacing of the bridges 12 of the first zone limit the pressure drops and avoid the entry of hot gases through the opening 17 of the cap 10.
  • FIGS 3, 4 and 5 there is shown the cap 10 of the blade which has flanges 18 defining with the upper face of the blade a groove 19 in which is engaged by sliding a plate 20 provided with calibrated orifices 21 , 22.
  • the plate 20, after positioning in the groove 19, can be fixed for example by soldering (shown in bold lines in Figure 4).
  • the diameter of the orifices 21, 22 can be optimized at will. Optimization of the diameter of the orifices 21 and 22, which are dusting holes, will be carried out by successive approaches during the tests, by changing the insert 20. For example, in a cooled blade where the cooling holes drilled in the leading edge and the trailing edge would have a diameter of 0.5 mm, it was found that a diameter of 1 mm was suitable for the orifices 21 and 22. It is even possible to completely seal the orifices 21 and 22 of the upstream 1 or downstream 2 cavities, or both, which makes it possible to increase the pressures in the cavities and to increase the flow rate near the leading edge or towards the trailing edge.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP19810400179 1980-02-19 1981-02-05 Perfectionnement aux aubes de turbines refroidies Expired EP0034961B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8003552A FR2476207A1 (fr) 1980-02-19 1980-02-19 Perfectionnement aux aubes de turbines refroidies
FR8003552 1980-02-19

Publications (2)

Publication Number Publication Date
EP0034961A1 EP0034961A1 (fr) 1981-09-02
EP0034961B1 true EP0034961B1 (fr) 1984-10-03

Family

ID=9238706

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19810400179 Expired EP0034961B1 (fr) 1980-02-19 1981-02-05 Perfectionnement aux aubes de turbines refroidies

Country Status (4)

Country Link
EP (1) EP0034961B1 (enrdf_load_stackoverflow)
JP (1) JPS56159507A (enrdf_load_stackoverflow)
DE (1) DE3166389D1 (enrdf_load_stackoverflow)
FR (1) FR2476207A1 (enrdf_load_stackoverflow)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11525360B2 (en) 2017-02-07 2022-12-13 Safran Helicopter Engines Ventilated high pressure blade of a helicopter turbine comprising an upstream duct and a central cooling chamber

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US4515526A (en) * 1981-12-28 1985-05-07 United Technologies Corporation Coolable airfoil for a rotary machine
US4474532A (en) * 1981-12-28 1984-10-02 United Technologies Corporation Coolable airfoil for a rotary machine
JPS5997205U (ja) * 1982-12-21 1984-07-02 株式会社東芝 ガスタ−ビンの翼
US4515523A (en) * 1983-10-28 1985-05-07 Westinghouse Electric Corp. Cooling arrangement for airfoil stator vane trailing edge
GB2152150A (en) * 1983-12-27 1985-07-31 Gen Electric Anti-icing inlet guide vane
US4650394A (en) * 1984-11-13 1987-03-17 United Technologies Corporation Coolable seal assembly for a gas turbine engine
JPS62271902A (ja) * 1986-01-20 1987-11-26 Hitachi Ltd ガスタ−ビン冷却翼
US4820123A (en) * 1988-04-25 1989-04-11 United Technologies Corporation Dirt removal means for air cooled blades
US4820122A (en) * 1988-04-25 1989-04-11 United Technologies Corporation Dirt removal means for air cooled blades
US4962640A (en) * 1989-02-06 1990-10-16 Westinghouse Electric Corp. Apparatus and method for cooling a gas turbine vane
FR2798421B1 (fr) * 1990-01-24 2002-10-11 United Technologies Corp Pales refroidies pour moteurs a turbine a gaz
FR2798423B1 (fr) * 1990-01-24 2002-10-11 United Technologies Corp Commande de jeu pour turbine de moteur a turbine a gaz
US5488825A (en) * 1994-10-31 1996-02-06 Westinghouse Electric Corporation Gas turbine vane with enhanced cooling
US5645397A (en) * 1995-10-10 1997-07-08 United Technologies Corporation Turbine vane assembly with multiple passage cooled vanes
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
US5842829A (en) * 1996-09-26 1998-12-01 General Electric Co. Cooling circuits for trailing edge cavities in airfoils
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US6406260B1 (en) 1999-10-22 2002-06-18 Pratt & Whitney Canada Corp. Heat transfer promotion structure for internally convectively cooled airfoils
US6257831B1 (en) 1999-10-22 2001-07-10 Pratt & Whitney Canada Corp. Cast airfoil structure with openings which do not require plugging
EP1167689A1 (de) * 2000-06-21 2002-01-02 Siemens Aktiengesellschaft Konfiguration einer kühlbaren Turbinenschaufel
US6609891B2 (en) * 2001-08-30 2003-08-26 General Electric Company Turbine airfoil for gas turbine engine
US6602047B1 (en) * 2002-02-28 2003-08-05 General Electric Company Methods and apparatus for cooling gas turbine nozzles
US6942449B2 (en) 2003-01-13 2005-09-13 United Technologies Corporation Trailing edge cooling
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US7021893B2 (en) * 2004-01-09 2006-04-04 United Technologies Corporation Fanned trailing edge teardrop array
US7001151B2 (en) * 2004-03-02 2006-02-21 General Electric Company Gas turbine bucket tip cap
US7165940B2 (en) * 2004-06-10 2007-01-23 General Electric Company Method and apparatus for cooling gas turbine rotor blades
US7520723B2 (en) 2006-07-07 2009-04-21 Siemens Energy, Inc. Turbine airfoil cooling system with near wall vortex cooling chambers
GB2441148A (en) * 2006-08-23 2008-02-27 Rolls Royce Plc Gas turbine engine component with coolant passages
US7607891B2 (en) * 2006-10-23 2009-10-27 United Technologies Corporation Turbine component with tip flagged pedestal cooling
US7934906B2 (en) 2007-11-14 2011-05-03 Siemens Energy, Inc. Turbine blade tip cooling system
FR2924155B1 (fr) * 2007-11-26 2014-02-14 Snecma Aube de turbomachine
FR2924156B1 (fr) * 2007-11-26 2014-02-14 Snecma Aube de turbomachine
FR2928405B1 (fr) * 2008-03-05 2011-01-21 Snecma Refroidissement de l'extremite d'une aube.
FR2954798B1 (fr) * 2009-12-31 2012-03-30 Snecma Aube a ventilation interieure
US20130052036A1 (en) * 2011-08-30 2013-02-28 General Electric Company Pin-fin array
US8790084B2 (en) * 2011-10-31 2014-07-29 General Electric Company Airfoil and method of fabricating the same
US20140064983A1 (en) * 2012-08-31 2014-03-06 General Electric Company Airfoil and method for manufacturing an airfoil
EP2832956A1 (de) * 2013-07-29 2015-02-04 Siemens Aktiengesellschaft Turbinenschaufel mit tragflächenprofilförmigen Kühlkörpern
EP3123000B1 (en) * 2014-03-27 2019-02-06 Siemens Aktiengesellschaft Blade for a gas turbine and method of cooling the blade
US10156146B2 (en) * 2016-04-25 2018-12-18 General Electric Company Airfoil with variable slot decoupling
GB201610783D0 (en) * 2016-06-21 2016-08-03 Rolls Royce Plc Trailing edge ejection cooling
EP3354850A1 (en) 2017-01-31 2018-08-01 Siemens Aktiengesellschaft A turbine blade or a turbine vane for a gas turbine
JP6345319B1 (ja) * 2017-07-07 2018-06-20 三菱日立パワーシステムズ株式会社 タービン翼及びガスタービン
CN112177685A (zh) * 2020-10-21 2021-01-05 中国航发沈阳发动机研究所 一种高压涡轮转子叶片尾缝冷却结构
CN112392550B (zh) * 2020-11-17 2021-09-28 上海交通大学 涡轮叶片尾缘针肋冷却结构及冷却方法、涡轮叶片

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11525360B2 (en) 2017-02-07 2022-12-13 Safran Helicopter Engines Ventilated high pressure blade of a helicopter turbine comprising an upstream duct and a central cooling chamber

Also Published As

Publication number Publication date
JPS6326242B2 (enrdf_load_stackoverflow) 1988-05-28
FR2476207B1 (enrdf_load_stackoverflow) 1983-05-13
FR2476207A1 (fr) 1981-08-21
DE3166389D1 (en) 1984-11-08
EP0034961A1 (fr) 1981-09-02
JPS56159507A (en) 1981-12-08

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