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EP1760261B1 - Umlenkeinrichtung der Luft des Kühlluftkreislaufs von Turbinenschaufeln - Google Patents

Umlenkeinrichtung der Luft des Kühlluftkreislaufs von Turbinenschaufeln Download PDF

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Publication number
EP1760261B1
EP1760261B1 EP06119232A EP06119232A EP1760261B1 EP 1760261 B1 EP1760261 B1 EP 1760261B1 EP 06119232 A EP06119232 A EP 06119232A EP 06119232 A EP06119232 A EP 06119232A EP 1760261 B1 EP1760261 B1 EP 1760261B1
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EP
European Patent Office
Prior art keywords
cavity
blade
cooling circuit
air
air deflector
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.)
Active
Application number
EP06119232A
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English (en)
French (fr)
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EP1760261A1 (de
Inventor
Jacques Auguste Boury
Patrice Eneau
Guy Moreau
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
SNECMA SAS
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Publication date
Application filed by SNECMA SAS filed Critical SNECMA SAS
Publication of EP1760261A1 publication Critical patent/EP1760261A1/de
Application granted granted Critical
Publication of EP1760261B1 publication Critical patent/EP1760261B1/de
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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
    • 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 the general field of cooling gas turbine blades, in particular the blades of a turbomachine gas turbine.
  • the gas turbine blades of a turbomachine such as the blades of the high-pressure turbine for example, are subjected to the very high temperatures of the gases from the combustion chamber. These temperatures reach values much higher than those that can withstand without damage the vanes of the turbine, which has the effect of limiting their service life.
  • a gas turbine blade In terms of mechanical strength, a gas turbine blade has a good service life if its intrados and extrados faces have similar temperatures (that is to say if the thermal gradient between these faces is low). Moreover, whatever the embodiment of the cooling circuits, the internal cooling of a turbine blade is provided by internal convection of a fresh air flow on the walls of the cavities forming these circuits. This results in a different heat exchange on each wall of the cavity, regardless of whether it is smooth or disturbed or that the blade is fixed or movable.
  • the heat exchange with the hot gases flowing outside the blade is greater on the intrados side than on the extrados side of the blade. Also, to compensate for this phenomenon and thus obtain a low thermal gradient between the intrados and extrados surfaces of the blade, it is necessary to strongly cool the internal walls of the cooling system cavities which are arranged on the intrados side of the blade.
  • the main object of the present invention is thus to overcome such drawbacks by proposing a gas turbine blade for which the internal cooling circuit makes it possible to minimize the difference in temperature between the intrados and extrados faces thereof.
  • a gas turbine blade according to claim 1 is provided.
  • the air deflector has a slope inclined so as to project the air flowing along the wall of the cavity to the opposite wall.
  • a ramp has a length of between 2 and 4 times its height and may have a radius of curvature of between 20 and 30 mm.
  • the wall of the cavity of the cooling circuit comprising the air deflector can be arranged on the extrados side of the blade and the wall of the cavity on which the air is projected can be arranged on the intrados side of the dawn.
  • the air deflector is advantageously disposed on the wall of the cavity of the cooling circuit at a zone of attachment of the blade.
  • the air deflector is advantageously disposed on the wall of the cavity of the cooling circuit at the top of the blade.
  • the air deflector may be positioned at a passage communicating the radial end of one of the cavities with a radial end adjacent to the cavity. other cavity.
  • the invention also relates to a gas turbine and a turbomachine having a plurality of blades as defined above.
  • the Figures 1 to 4 represent a moving blade 10 of a turbomachine, such as a moving blade of a high-pressure turbine.
  • a turbomachine such as a moving blade of a high-pressure turbine.
  • the invention can be applied to other blades of a turbomachine gas turbine as well as to vanes of a turbomachine gas turbine.
  • the blade 10 comprises an aerodynamic surface (or blade) which extends radially between a blade root 12 and a blade tip 14.
  • This aerodynamic surface consists of a leading edge 16 disposed opposite the blade. flow of the hot gases from the combustion chamber of the turbomachine, a trailing edge 18 opposite the leading edge 16, a lateral face 20 and an extrados lateral face 22, these lateral faces 20 , 22 connecting the leading edge 16 to the trailing edge 18.
  • the blade 10 is provided with an internal cooling circuit of the type formed by at least one cavity extending radially between the root 12 and the top 14 of the blade, at least one air intake opening at a radial end of the cavity and at least one air outlet opening opening into the cavity and opening on one of the faces of the blade.
  • the internal cooling circuit of the blade consists of a leading edge cavity 24 disposed on the side of the leading edge 16 of the blade, three central cavities 26, 28 and 30 disposed in a central portion of the dawn and a trailing edge cavity 32 disposed on the side of the trailing edge 18 of the blade.
  • These different cavities 24, 26, 28, 30 and 32 extend from the intrados face 20 to the extrados face 22 of the blade.
  • An air inlet opening 34 is provided at a radial end of the leading edge cavity 24 (here at the foot 12 of the blade) to supply air to the cooling circuit.
  • a first passage 36 communicates the other radial end of the leading edge cavity 24 with a radial end adjacent to the central cavity 26 adjacent.
  • a second passage 38 and a third passage 40 communicate respectively the central cavity 26 with the central cavity 28 adjacent and the latter with the central cavity 30 remaining.
  • a fourth passage 42 communicates the central cavity 30 with the trailing edge cavity 32.
  • the intrados cooling circuit also has outlet orifices 44 opening in the trailing edge cavity 32 and opening on the intrados face 20 of the blade at the trailing edge 18 of the latter. These orifices 44 are regularly distributed over the entire radial height of the blade.
  • Airflow disturbers 46 for increasing heat transfer may be provided along the walls of the various cavities 24, 26, 28, 30 and 32 of the cooling circuit. These flow disturbers 46 may be in the form of ribs which are straight or inclined relative to the axis of rotation of the blade, in the form of pins or in any other equivalent form.
  • any other embodiment of the internal cooling circuit of the blade of the type described above is applicable to the invention.
  • the number, shape and arrangement of the cavities, as well as the quantity and arrangement of the air intake orifices, the communication passages and the outlet orifices may vary according to the cooling circuit.
  • At least one of the walls of one (or more) of the cavities 24, 26, 28, 30 and 32 of the cooling circuit comprises at least one air deflector 48, 48 '.
  • an air deflector 48 is particularly visible on the figures 2 and 3 .
  • the air deflector 48 is positioned on the wall 24a of the leading edge cavity 24 which is disposed on the extrados side 22 of the blade.
  • FIG. 4 Another example of the location of such an air deflector 48 'is shown on the figure 4 .
  • the air deflector 48 ' is disposed on the wall 26a of the central cavity 26 adjacent to the leading edge cavity 24 which is disposed on the extrados side 22 of the blade.
  • the shape and the dimensions of the air deflector 48, 48 ' are adapted to project the air flowing along the wall 24a, 26a of the cavity 24, 26 to an opposite wall 24b, 26b of the cavity while avoiding a re-bonding of the boundary layer immediately downstream of the air deflector.
  • the air deflector according to the invention is distinguished in that it consists, on the one hand to project the air on the wall opposite to that of its implantation, and secondly to avoid an immediate gluing of the boundary layer.
  • an airflow disturbance has the essential function of increasing the turbulence of the flow of air in the immediate vicinity of the disturbance while seeking to re-glue the flow downstream thereof.
  • the presence of airflow disruptors 46 with the air deflector 48, 48 'according to the invention is also not incompatible.
  • the figure 3 shows more precisely one embodiment of an air deflector 48 according to the invention.
  • the air deflector 48 comprises a ramp 52 which is inclined with respect to the wall 24a of the cavity 24 on which the deflector is implanted so as to project the air flowing along this wall 24a towards the opposite wall 24b .
  • the inclined ramp 52 of the air deflector 48 has a length L which is between 2 and 4 times its height h .
  • the ramp 52 of the air deflector 48 has a height h of the order of 1.5 mm and a length L of between 3 and 5 mm.
  • an air-flow disruptor 46 as described above has a height of between 0.4 and 0.5 mm. .
  • the inclined ramp 52 of the air deflector 48 is rounded and has a radius of curvature R between 20 and 30 mm.
  • This value is given as an example for a cooling cavity 24 having a width d of the order of 4 mm.
  • a radius of curvature R as large relative to the width d of the cavity 24 makes it possible to move the air flowing along the wall 24a towards the opposite wall 24b without accelerating it suddenly.
  • the radius of curvature R of the ramp 52 of the deflector is preferably greater than the length L on which extends this ramp.
  • the air deflector 48 On the opposite side to the inclined ramp 52, the air deflector 48 has another rounded ramp 54 whose radius of curvature r and the length I on which it extends are calculated so as to avoid re-bonding of the boundary layer immediately. downstream of the air deflector.
  • the radius of curvature r of this other ramp 54 must be as small as possible to achieve this goal.
  • the flow of air in the leading edge cavity 24 is centrifugal, that is to say that the air flows from the foot 12 to the top 14 of the blade.
  • the air deflector 48 is advantageously disposed on the wall of the cavity 24 of the cooling circuit at a zone of attachment of the blade. This attachment zone extends from the radial end of the blade on the side of its foot 12 to a platform 56 defining the inner wall of the flow passage of the gas passing through the gas turbine.
  • a platform 56 defining the inner wall of the flow passage of the gas passing through the gas turbine.
  • the flow of air in the central cavity 26 is centripetal, that is to say that the air flows from the top 14 towards the foot 12 of dawn.
  • the air deflector 48 ' is advantageously disposed on the wall of the cavity 26 of the cooling circuit at the top 14 of the blade. Such a location makes it possible to obtain an optimum internal heat exchange on the underside of the dawn.
  • the air deflector 48 is positioned at a passage 100 communicating the radial end of a cavity 102 of an internal cooling circuit of a blade with a radial end adjacent to another cavity 104 adjacent thereto, such a communication passage 100 may for example be one of the passages 36 to 40 of the dawn of Figures 1 to 3 .
  • the air baffle 48 is disposed on one of the walls 104a of the cavity 104 and its shape and dimensions are adapted to project the air flowing along the wall 104a to the opposite wall 104b while avoiding a bonding of the boundary layer immediately downstream of the air deflector.
  • the air baffle 48 is positioned such that the air flowing in the cavity 102 is projected at its" turn-over "into the adjacent cavity 104 (i.e. from the communication passage 100) to an air circulation zone 106 which is located at the radial end of the opposite wall 104b of the adjacent cavity 104.
  • Such an area 106 is usually an area in which the Air circulation is low and undisturbed.
  • the air deflector 48 "thus makes it possible to avoid any risk of detachment of the boundary layer at the level of the" reversal "zone of the air between the two cavities 102, 104 of the cooling circuit .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (9)

  1. Gasturbinenschaufel (10) umfassend einen inneren Kühlkreis, der aus wenigstens einem sich radial zwischen dem Fuß (12) und der Spitze (14) der Schaufel erstreckenden Hohlraum (24, 26, 102, 104) besteht, wenigstens eine Luftzufuhröffnung (34) an einem radialen Ende des Hohlraums (24, 26, 102, 104) und wenigstens eine Luftaustrittsöffnung (44), die sich in den Hohlraum öffnet und an einer der Seiten (20, 22) der Schaufel ausmündet, wobei wenigstens eine der Wände (24a, 26a, 104a) des Hohlraums des Kühlkreises wenigstens eine Luftablenkeinrichtung (48, 48', 48") aufweist, deren Form und Abmessungen entsprechend ausgeführt sind, um die entlang der Wand (24a, 26a, 104a) des Hohlraums strömende Luft in Richtung einer gegenüberliegenden Wand (24b, 26b, 104b) des Hohlraums zu lenken und dabei ein Wiederanlegen der Grenzschicht unmittelbar stromabwärts der Luftablenkeinrichtung (48, 48', 48") zu vermeiden, dadurch gekennzeichnet, daß die Luftablenkeinrichtung (48, 48', 48") eine geneigte Rampe (52) aufweist, die abgerundet ist und die eine Länge (L) aufweist, welche zwei- bis viermal so groß ist wie ihre Höhe (h), so daß die entlang der Wand (24a, 26a, 104a) des Hohlraums strömende Luft in Richtung der gegenüberliegenden Wand (24b, 26b, 104b) gelenkt wird, wobei die geneigte Rampe (52) der Luftablenkeinrichtung eine Höhe (h) aufweist, die etwa 37,5 % des Abstands zwischen den zwei gegenüberliegenden Wänden des Hohlraums des Kühlkreises entspricht.
  2. Schaufel nach Anspruch 1, bei der die geneigte Rampe (52) der Luftablenkeinrichtung (48, 48', 48") einen Krümmungsradius (R) zwischen 20 und 30 mm aufweist.
  3. Schaufel nach einem der Ansprüche 1 und 2, bei der die die Luftablenkeinrichtung (48, 48') aufweisende Wand (24a, 26a) des Hohlraums (24, 26) des Kühlkreises auf der Rückseite (22) der Schaufel angeordnet ist und die Wand (24b, 26b) des Hohlraums, auf die die Luft gelenkt wird, auf der Vorderseite (20) der Schaufel angeordnet ist.
  4. Schaufel nach einem der Ansprüche 1 bis 3, bei der die Luftablenkeinrichtung (48) an der Wand (24a) des Hohlraums (24) des Kühlkreises in Höhe eines Befestigungsbereichs der Schaufel angeordnet ist.
  5. Schaufel nach einem der Ansprüche 1 bis 4, bei der die Luftablenkeinrichtung (48') an der Wand (26a) des Hohlraums (26) des Kühlkreises im Bereich der Spitze (14) der Schaufel angeordnet ist.
  6. Schaufel nach einem der Ansprüche 1 bis 3, bei welcher der innere Kühlkreis wenigstens zwei Hohlräume (102, 104) umfaßt, wobei die Luftablenkeinrichtung (48") im Bereich eines Durchgangs (100) angeordnet ist, der das radiale Ende eines Hohlraums (102) mit einem benachbarten radialen Ende des anderen Hohlraums (104) verbindet.
  7. Schaufel nach einem der Ansprüche 1 bis 6, bei der die Wände des Hohlraums (24, 26) des Kühlkreises mit einer Vielzahl von Strömungsstörelementen (46) versehen sind, die dazu bestimmt sind, die Wärmeübertragungen entlang dieser Wände zu erhöhen.
  8. Gasturbine, die eine Vielzahl von Schaufeln nach einem der Ansprüche 1 bis 7 umfaßt.
  9. Turbomaschine mit einer Gasturbine, die eine Vielzahl von Schaufeln nach einem der Ansprüche 1 bis 7 aufweist.
EP06119232A 2005-08-25 2006-08-21 Umlenkeinrichtung der Luft des Kühlluftkreislaufs von Turbinenschaufeln Active EP1760261B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0508740A FR2890103A1 (fr) 2005-08-25 2005-08-25 Deflecteur d'air pour circuit de refroidissement pour aube de turbine a gaz

Publications (2)

Publication Number Publication Date
EP1760261A1 EP1760261A1 (de) 2007-03-07
EP1760261B1 true EP1760261B1 (de) 2008-03-05

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP06119232A Active EP1760261B1 (de) 2005-08-25 2006-08-21 Umlenkeinrichtung der Luft des Kühlluftkreislaufs von Turbinenschaufeln

Country Status (6)

Country Link
US (1) US7192251B1 (de)
EP (1) EP1760261B1 (de)
CA (1) CA2557112C (de)
DE (1) DE602006000641T2 (de)
ES (1) ES2303312T3 (de)
FR (1) FR2890103A1 (de)

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Publication number Priority date Publication date Assignee Title
GB0800361D0 (en) * 2008-01-10 2008-02-20 Rolls Royce Plc Blade cooling
US10830096B2 (en) 2013-10-03 2020-11-10 Raytheon Technologies Corporation Rotating turbine vane bearing cooling
US9551229B2 (en) 2013-12-26 2017-01-24 Siemens Aktiengesellschaft Turbine airfoil with an internal cooling system having trip strips with reduced pressure drop
US10184341B2 (en) 2015-08-12 2019-01-22 United Technologies Corporation Airfoil baffle with wedge region
US10012092B2 (en) 2015-08-12 2018-07-03 United Technologies Corporation Low turn loss baffle flow diverter
KR101797370B1 (ko) * 2016-07-04 2017-12-12 두산중공업 주식회사 가스터빈 블레이드
RU183316U1 (ru) * 2018-04-09 2018-09-18 Федеральное государственное бюджетное образовательное учреждение высшего образования "Рыбинский государственный авиационный технический университет имени П.А. Соловьева" Дефлектор охлаждаемой сопловой турбинной лопатки
US10774657B2 (en) 2018-11-23 2020-09-15 Raytheon Technologies Corporation Baffle assembly for gas turbine engine components
KR102161765B1 (ko) * 2019-02-22 2020-10-05 두산중공업 주식회사 터빈용 에어포일, 이를 포함하는 터빈
FR3107920B1 (fr) 2020-03-03 2023-11-10 Safran Aircraft Engines Aube creuse de turbomachine et plateforme inter-aubes équipées de saillies perturbatrices de flux de refroidissement
CN113550794B (zh) * 2021-09-10 2022-12-06 中国航发湖南动力机械研究所 一种涡轮转子叶片的多腔高效冷却结构及其冷却方法

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US3171631A (en) * 1962-12-05 1965-03-02 Gen Motors Corp Turbine blade
US4180373A (en) * 1977-12-28 1979-12-25 United Technologies Corporation Turbine blade
US4775296A (en) * 1981-12-28 1988-10-04 United Technologies Corporation Coolable airfoil for a rotary machine
DE19526917A1 (de) * 1995-07-22 1997-01-23 Fiebig Martin Prof Dr Ing Längswirbelerzeugende Rauhigkeitselemente
DE19612840A1 (de) * 1996-03-30 1997-10-02 Abb Research Ltd Vorrichtung und Verfahren zur Kühlung einer einseitig von Heissgas umgebenen Wand
US5738493A (en) * 1997-01-03 1998-04-14 General Electric Company Turbulator configuration for cooling passages of an airfoil in a gas turbine engine
US6154571A (en) * 1998-06-24 2000-11-28 Nec Research Institute, Inc. Robust digital watermarking
FR2829175B1 (fr) * 2001-08-28 2003-11-07 Snecma Moteurs Circuits de refroidissement pour aube de turbine a gaz
FR2858352B1 (fr) * 2003-08-01 2006-01-20 Snecma Moteurs Circuit de refroidissement pour aube de turbine
GB2405451B (en) * 2003-08-23 2008-03-19 Rolls Royce Plc Vane apparatus for a gas turbine engine

Also Published As

Publication number Publication date
DE602006000641D1 (de) 2008-04-17
FR2890103A1 (fr) 2007-03-02
EP1760261A1 (de) 2007-03-07
US7192251B1 (en) 2007-03-20
CA2557112A1 (fr) 2007-02-25
ES2303312T3 (es) 2008-08-01
DE602006000641T2 (de) 2009-03-26
CA2557112C (fr) 2013-12-10
US20070048136A1 (en) 2007-03-01

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