EP1557533A1 - Refroidissement d'une aube de turbine avec faux-plancher entre l'aube et son extrémité - Google Patents

Refroidissement d'une aube de turbine avec faux-plancher entre l'aube et son extrémité Download PDF

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Publication number
EP1557533A1
EP1557533A1 EP04001468A EP04001468A EP1557533A1 EP 1557533 A1 EP1557533 A1 EP 1557533A1 EP 04001468 A EP04001468 A EP 04001468A EP 04001468 A EP04001468 A EP 04001468A EP 1557533 A1 EP1557533 A1 EP 1557533A1
Authority
EP
European Patent Office
Prior art keywords
blade
wall
turbine
cooling
turbine blade
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
EP04001468A
Other languages
German (de)
English (en)
Other versions
EP1557533B1 (fr
Inventor
Stefan Dr. Baldauf
Hans-Thomas Dr. Bolms
Michael HÄNDLER
Christian Lerner
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.)
Siemens AG
Siemens Corp
Original Assignee
Siemens AG
Siemens Corp
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 Siemens AG, Siemens Corp filed Critical Siemens AG
Priority to DE200450006484 priority Critical patent/DE502004006484D1/de
Priority to EP20040001468 priority patent/EP1557533B1/fr
Publication of EP1557533A1 publication Critical patent/EP1557533A1/fr
Application granted granted Critical
Publication of EP1557533B1 publication Critical patent/EP1557533B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime 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
    • F01D5/186—Film 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
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by 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
    • F05D2260/221—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 invention relates to a turbine blade with a for Tip of the turbine blade, along a blade axis, arranged in the form of a hollow profile airfoil and with a blade tip in the form of a hollow profile.
  • the invention also relates to a gas turbine with one along an axis the gas turbine extending flow channel with annular Cross section for acting on a working fluid and a number of blade stages arranged along the axis, wherein a blade stage is a number of annular arranged radially in the flow channel extending Having turbine blades, which can be acted upon by a cooling fluid are.
  • temperatures occur in the flow channel on, which lie in the range between 1000 ° C and 1400 ° C if this with a working fluid in the form of Hot gas is applied.
  • Turbine blades for recording the kinetic energy of the working fluid same are exposed to such pressures interpreted.
  • a turbine blade can take the form of a a rotor and attached to the rotor blade formed his.
  • both types of blades especially the tip of the turbine blade high thermal Exposed to stress. When a blade occurs In addition, the occurring due to the rotational movement of high mechanical Add burden.
  • the outer shape of the blade tip is primarily through an aerodynamic task determined. Farther play in the shaping manufacturing reasons one essential role.
  • a cooling technology Interpretation relevant. Namely long service life To reach, the tip must be cooled. Without Cooling the blade tip would rapidly oxidize it. The Lifetime requirements for vanes, and in particular on blades, but are still getting bigger. That means that Component temperature must be through the use of a cooling fluid be kept within reasonable limits.
  • the use of cooling fluid can be made efficient to the Overall efficiency and to increase the performance of a gas turbine.
  • a turbine blade and in particular a Turbine blade tip are proposed a number of measures Service.
  • the US 4,519,745 provides for cooling a Blade in the form of a hollow profile a corrugated Distribution of the inner wall of the hollow profile before.
  • a Blade tip is a variety of channels and one Recess in a bearing surface for the blade tip provided so that the blade tip from the outside with one transported over the channels and the recess Cooling fluid can be cooled.
  • the invention begins, whose task it is a turbine blade and a gas turbine with a turbine blade indicate that improved cooling of the Blade tip is provided.
  • the task is accomplished by the Invention solved by means of the turbine blade mentioned, in the present invention, the airfoil at his the Turbine blade end facing one in the hollow profile has transverse to the blade axis extending first wall and the blade tip at its remote from the turbine blade end Side one in the hollow profile transverse to the blade axis extending second wall, wherein the first wall and the second wall to form a cooling Opposite supporting floor.
  • the cooling means are in particular for cooling provided the blade tip and represent measures that an improved heat transfer between the blade and the cooling medium support.
  • the essential realization of the invention is that the one formed by the first and the second wall Double bottom particularly effective for attaching Cooling agents as the heat transfer improving Measure is appropriate and thus to a particularly efficient Cooling of the blade tip can be used.
  • the Invention is based on the consideration that the geometry a double floor a particularly suitable basis for a cooling agent. There are different ones Types of cooling equipment in question, all within the framework of the geometry of said double floor are very effectively interpretable.
  • the first and / or the second are Wall around a wall that extends over the entire cross section of the hollow profile in the interior of the hollow profile transverse to the blade axis extends, preferably perpendicular to the blade axis, so horizontal.
  • the formed according to the concept of the invention and cooling means carrying double bottom sees So the formation of a cavity between the first wall and the second wall in front.
  • the cooling means is a means from the group consisting of: turbulator agent, impingement coolant and film coolant. That come useful to Increase of cooling efficiency the principle of turbulators, Impact cooling and film cooling or other measures to Increase of the cooling effect of the used cooling medium or for lowering the external hot gas temperature individually or in Combination for use.
  • the turbulator contributes to the formation of the turbulator means the first wall and / or the second wall a number of swirling elements.
  • Such swirling serve for Turbulation of a cooling fluid when the turbine blade with a cooling fluid is applied.
  • This is in particular the Turbulence of the cooling fluid between the first wall and the second wall, ie in the above-mentioned cavity, appropriate.
  • Suitably serves an arrangement of nipples, dimples or other suitable formations, e.g. Stege u.a., on the hollow profile or the walls.
  • Further turbulence elements also have a beneficial effect in terms of heat dissipation from the blade tip, such as a suitable one Arrangement of ribs or joints.
  • the first wall has a number of impingement cooling holes that allow it when the turbine blade is exposed to cooling fluid, bouncing the cooling fluid onto the second wall.
  • An impingement cooling opening in the first Wand is characterized in particular by the fact that they are in one Angle in the range of about 0 ° measured to the blade axis, that is perpendicular to the second wall, is arranged.
  • a fair spatial proximity of the second wall lead to the first wall to that one through the baffle cooling holes with suitable Compressive cooling fluid bounces on the second wall and
  • a film coolant the second wall and / or the hollow profile of the blade tip contributes to the formation of a film coolant the second wall and / or the hollow profile of the blade tip a Number of film cooling holes that allow it to be applied the turbine blade with cooling fluid a film of cooling fluid to create on the wall structure of the blade tip.
  • film cooling openings thus lead in the direction the turbine blade end obliquely from inside to outside and conduct cooling fluid to the inner wall and / or outer wall of the Blade tip.
  • a film cooling opening is thus essentially measured at an angle of ⁇ 90 ° and significantly greater than 0 ° arranged to the blade axis in the second wall.
  • film cooling holes for cooling the airfoil possibly also on the first wall, be provided. Further Measures can be through additional cooling holes or cooling channels will be realized.
  • the wall structure of the blade tip a higher porosity and / or have a smaller wall thickness than the wall structure of the airfoil.
  • the second wall is a closure means for a core holding hole in the first wall.
  • a retaining bore for the blade also referred to as core holding bore provided.
  • the second wall is a plug or other closure means for the core holding bore, so in one assembled condition of airfoil and Blade tip the closure means on the second wall the Core holding hole in the first wall closes.
  • the airfoil and the blade tip from different Materials are poured.
  • that is Airfoil cast from a high-strength material.
  • the blade tip is made of a highly heat-conductive Poured material.
  • This type of training is the consideration underlying that the airfoil especially in terms of mechanical requirements, mainly due to the rotational Load a blade should be designed.
  • especially the blade tip in terms of cooling technology Measures due to one over one Blade sheet higher thermal load interpret.
  • the blade tip is advantageous to the blade in the Soldered and / or welded frame of a joining process.
  • a mechanical connection is also additional or alternative possible.
  • the invention also leads to a gas turbine initially mentioned Type in which the turbine blades explained above Art are formed.
  • Figure 1 shows a on a rotor 3 a not shown Gas Turbine Attached Blade 1.
  • the Blade 1 is one of a number of annularly arranged, radially into a flow channel 5 of the gas turbine extending turbine blades, which in their entirety a Shovel stage form, resulting in the annular cross-section the flow channel 5 extends.
  • a variety of such annular blade stages is as well as the Flow channel along an axis 7 of the not shown Gas turbine arranged.
  • the flow channel 5 is with a working fluid 9 in the form of a hot gas mixture applied, which relaxes under the drive of the blade 1 and thus its kinetic energy with rotation of the rotor to drive a generator, not shown.
  • Blade 1 points toward turbine blade end 29, arranged one after the other along their blade axis 11
  • Platform area 13 arranged in the form of a hollow profile Airfoil 15 and arranged in the form of a hollow profile Blade tip 17 on.
  • the platform area 13 comprises while a paddle platform for limiting the Flow channel 5 and a blade root, in detail are not shown.
  • the blade 1 is as present indicated schematically, via a channel system 19 with a Cooling fluid 21 acted upon.
  • the cooling system 19 has also via suitable dosing means 23, the feeder 25th of the cooling fluid 21 in the blade can control and the indicated here only schematically by the reference numeral 23 are.
  • the cooling of the blade 1 extends in particular also on the blade 15 and the blade tip 17th
  • FIGS. 2 and 3 a first and second are shown particularly preferred embodiment of the design of the cooling system 19 in the area 27 between the blade 15 and the blade tip 17 shown.
  • the transition area 27 in particular includes the blade 15 at its the Turbine blade end 29 of the turbine blade 1 facing Side and the blade tip 17 at its from the turbine blade end 29 of the blade 1 side facing away.
  • FIG 2 shows a first preferred embodiment of a Blade 31 in the transition region 27 of Figure 1 in a perspective sectional view.
  • the blade 31 has a trained as a hollow profile airfoil 33 and a formed as a hollow profile blade tip 35.
  • the hollow profile of the blade 33 has a cavity 37 on and the hollow profile of the blade tip 35 has a Cavity 39, which is part of the illustrated in Figure 1 Cooling system 19 and can be acted upon with cooling fluid.
  • the airfoil 33 has at its one shown in Figure 1 Turbine blade end 29 of the turbine blade 31 facing Side transverse to the blade axis 11 extending first wall 41.
  • the blade tip 35 has at its from the turbine blade end 29 of the turbine blade shown in Figure 1 1 opposite side transverse to the blade axis 11 extending second wall 43.
  • the first wall 41 and the second wall 43 are thereby forming a double floor 45 opposite.
  • a double bottom 45 is between the first wall 41 and the second wall 43, the both horizontally over the entire cross section of the hollow profile extend, a cavity 47 formed.
  • the raised floor 45 carries thereby cooling technical means, which in the following are explained in detail.
  • the second carries Wall 43 a number of vortex or turbulator elements in the form of nipples 49 and dimples 51.
  • a nipple 49 is thereby formed in the cavity 47 into the second wall 43.
  • a dimple 51 extends the cavity 47 in the form of a Recess in the second wall 43.
  • the aforementioned swirling elements mainly serve to swirl a cooling fluid, through the impingement cooling holes 53 in the first Wall 41 can be supplied to the cavity 47.
  • the impact cooling openings 53 are perpendicular to the second wall 43rd arranged.
  • first wall 41 is so close the second wall 43 arranged that a corresponding pressurized Cooling medium on the impingement cooling openings 53 on the second wall 43 bounces and the blade tip 35 over the second web 43 effectively cools in the context of an impingement cooling.
  • swirling elements in the form of nipples 49 and dimpling 51 is this explained impact cooling yet strengthened. That is, heat absorbed in the blade tip is thereby effectively dissipated by the cooling fluid.
  • the second wall 43 has a first number of Film cooling holes 55 and a second number of film cooling holes 57 on.
  • a first film cooling opening 55 is in Direction of the turbine blade end 29 of the turbine blade 31st oriented obliquely from the inside out and can thus open up adequately gentle way a film of cooling fluid on the inner wall 59 make the blade tip 35.
  • An even more obliquely arranged second film cooling opening 57 allows it's about the same mechanism another movie from cooling fluid on the outer wall 61 of the blade tip 35 produce.
  • the wall structure of the blade tip 35th a wall thickness 63 which is less than the wall thickness 65 of the wall structure of the airfoil 33.
  • a higher porosity not shown, than the wall structure of the airfoil 33.
  • first embodiment of a blade 31 is thus a cavity 47 comprising double floor 45 of the blade 31 advantageously with cooling technology Provided means that are better than before allow the blade tip 35 to effectively cool and with it Save cooling medium, which is the overall efficiency of a Gas turbine can increase.
  • a blade 31 advantageously in the context of a manufacturing process producing a separate casting process, and thus a separate production of the airfoil 33rd on the one hand and the blade tip on the other hand provides. The separately cast blade tip 35 is then disconnected cast airfoil 33 to form the double bottom 45 and the cavity 47 soldered.
  • the most advantageous Raised floor 45 with cavity 47 and Cooling means can therefore be particularly useful via a separate pouring process for blade tip 35 and airfoil 33 reach as explained.
  • the airfoil 33 from another Material can be poured as the blade tip 35.
  • the airfoil 33 is expediently in terms its greater mechanical stress from one high strength material poured while the blade tip 35 in terms of their higher thermal stress a high thermal conductivity material, e.g. a cobalt material, is poured.
  • FIG. 1 A particularly preferred second embodiment of a blade 1 is shown in FIG to a simplification of the manufacturing process in addition to the first embodiment shown in Figure 2 has further elements.
  • the remaining features of Figure 2 are provided in Figure 3 with the same reference numerals.
  • the one shown here Core holding hole 77 which is available by casting, be closed particularly reliable.
  • the connection between blade tip 85 and blade 73 is present by a particularly suitable welding method in the joining area 87 produced.
  • the connecting surfaces between Leave bucket tip 85 and blade 73 in joint area 87 Produce very easily and precisely.
  • the closure the core holding hole 77 is in the context of a special measured fit of the plug 81 on the Core holding hole 77.
  • Both in the first preferred embodiment shown in FIG the blade 31 as well as in the in FIG. 3 shown particularly preferred second embodiment of Blade 71 can be in the context of a separate production of airfoil 33, 73 and blade tip 35, 85 nevertheless relatively complex cooling means in a relatively simple way as part of a casting Provide manufacturing process in the blade profile 31, 71 and thus a particularly efficient cooling of the blade tip Reach 35, 85.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP20040001468 2004-01-23 2004-01-23 Refroidissement d'une aube de turbine avec faux-plancher entre l'aube et son extrémité Expired - Lifetime EP1557533B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
DE200450006484 DE502004006484D1 (de) 2004-01-23 2004-01-23 Kühlung einer Turbinenschaufel mit einem Doppelboden zwischen Schaufelblatt und Schaufelspitze
EP20040001468 EP1557533B1 (fr) 2004-01-23 2004-01-23 Refroidissement d'une aube de turbine avec faux-plancher entre l'aube et son extrémité

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP20040001468 EP1557533B1 (fr) 2004-01-23 2004-01-23 Refroidissement d'une aube de turbine avec faux-plancher entre l'aube et son extrémité

Publications (2)

Publication Number Publication Date
EP1557533A1 true EP1557533A1 (fr) 2005-07-27
EP1557533B1 EP1557533B1 (fr) 2008-03-12

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DE (1) DE502004006484D1 (fr)

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1614860A3 (fr) * 2004-07-08 2008-11-26 United Technologies Corporation Aube de turbine
EP2426316A1 (fr) * 2010-09-03 2012-03-07 Siemens Aktiengesellschaft Pale de turbine
FR2983517A1 (fr) * 2011-12-06 2013-06-07 Snecma Aube de turbine refroidie pour moteur a turbine a gaz.
EP2881541A1 (fr) * 2013-12-05 2015-06-10 Rolls-Royce Deutschland Ltd & Co KG Refroidissement de pointe d'aube d'un aube de rotor de turbine à gaz
EP3088673A1 (fr) * 2015-04-28 2016-11-02 Siemens Aktiengesellschaft Aube rotorique de turbine à gaz, rotor, turbine à gaz et moteur associés
EP3118414A1 (fr) * 2015-07-02 2017-01-18 United Technologies Corporation Surface portante de moteur à turbine à gaz
WO2019177600A1 (fr) * 2018-03-14 2019-09-19 General Electric Company Ensemble refroidissement pour ensemble turbine
EP3623577A1 (fr) * 2018-09-12 2020-03-18 United Technologies Corporation Agencement de refroidissement d'extrémité d'aube de turbine avec séparation de purge
CN114776387A (zh) * 2022-05-06 2022-07-22 中国联合重型燃气轮机技术有限公司 透平动叶和燃气轮机
CN115163203A (zh) * 2022-06-24 2022-10-11 中国船舶重工集团公司第七0三研究所 一种具有空心叶冠结构的涡轮动叶
CN121803307A (zh) * 2026-03-10 2026-04-07 中国航发沈阳发动机研究所 一种高压涡轮导向叶片双层壁结构

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2019177598A1 (fr) 2018-03-14 2019-09-19 General Electric Company Ensemble de refroidissement pour ensemble turbine

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3732031A (en) * 1970-06-17 1973-05-08 Gen Motors Corp Cooled airfoil
JPH02140402A (ja) * 1988-11-21 1990-05-30 Toshiba Corp ターボ機械の動翼構造
US4946346A (en) * 1987-09-25 1990-08-07 Kabushiki Kaisha Toshiba Gas turbine vane
GB2279705A (en) * 1985-07-24 1995-01-11 Rolls Royce Plc Cooling of turbine blades of a gas turbine engine
US5752802A (en) * 1996-12-19 1998-05-19 Solar Turbines Incorporated Sealing apparatus for airfoils of gas turbine engines
US6135715A (en) * 1999-07-29 2000-10-24 General Electric Company Tip insulated airfoil
US6164914A (en) * 1999-08-23 2000-12-26 General Electric Company Cool tip blade
US20030026698A1 (en) * 2001-08-02 2003-02-06 Flodman David Allen Trichannel airfoil leading edge cooling

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3732031A (en) * 1970-06-17 1973-05-08 Gen Motors Corp Cooled airfoil
GB2279705A (en) * 1985-07-24 1995-01-11 Rolls Royce Plc Cooling of turbine blades of a gas turbine engine
US4946346A (en) * 1987-09-25 1990-08-07 Kabushiki Kaisha Toshiba Gas turbine vane
JPH02140402A (ja) * 1988-11-21 1990-05-30 Toshiba Corp ターボ機械の動翼構造
US5752802A (en) * 1996-12-19 1998-05-19 Solar Turbines Incorporated Sealing apparatus for airfoils of gas turbine engines
US6135715A (en) * 1999-07-29 2000-10-24 General Electric Company Tip insulated airfoil
US6164914A (en) * 1999-08-23 2000-12-26 General Electric Company Cool tip blade
US20030026698A1 (en) * 2001-08-02 2003-02-06 Flodman David Allen Trichannel airfoil leading edge cooling

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
PATENT ABSTRACTS OF JAPAN *

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1614860A3 (fr) * 2004-07-08 2008-11-26 United Technologies Corporation Aube de turbine
EP2426316A1 (fr) * 2010-09-03 2012-03-07 Siemens Aktiengesellschaft Pale de turbine
WO2012028584A1 (fr) 2010-09-03 2012-03-08 Siemens Aktiengesellschaft Aube de turbine
FR2983517A1 (fr) * 2011-12-06 2013-06-07 Snecma Aube de turbine refroidie pour moteur a turbine a gaz.
GB2497420A (en) * 2011-12-06 2013-06-12 Snecma Turbine blade cooling
GB2497420B (en) * 2011-12-06 2016-04-13 Snecma Cooled turbine blade for gas turbine engine
US9435210B2 (en) 2011-12-06 2016-09-06 Snecma Cooled turbine blade for gas turbine engine
EP2881541A1 (fr) * 2013-12-05 2015-06-10 Rolls-Royce Deutschland Ltd & Co KG Refroidissement de pointe d'aube d'un aube de rotor de turbine à gaz
EP3088673A1 (fr) * 2015-04-28 2016-11-02 Siemens Aktiengesellschaft Aube rotorique de turbine à gaz, rotor, turbine à gaz et moteur associés
CN106089315A (zh) * 2015-04-28 2016-11-09 西门子股份公司 用于燃气涡轮的转子叶片
EP3118414A1 (fr) * 2015-07-02 2017-01-18 United Technologies Corporation Surface portante de moteur à turbine à gaz
US10053992B2 (en) 2015-07-02 2018-08-21 United Technologies Corporation Gas turbine engine airfoil squealer pocket cooling hole configuration
WO2019177600A1 (fr) * 2018-03-14 2019-09-19 General Electric Company Ensemble refroidissement pour ensemble turbine
JP2021520463A (ja) * 2018-03-14 2021-08-19 ゼネラル・エレクトリック・カンパニイ タービンアセンブリのため冷却アセンブリ
JP7150863B2 (ja) 2018-03-14 2022-10-11 ゼネラル・エレクトリック・カンパニイ タービンアセンブリのため冷却アセンブリ
EP3623577A1 (fr) * 2018-09-12 2020-03-18 United Technologies Corporation Agencement de refroidissement d'extrémité d'aube de turbine avec séparation de purge
US10801334B2 (en) 2018-09-12 2020-10-13 Raytheon Technologies Corporation Cooling arrangement with purge partition
CN114776387A (zh) * 2022-05-06 2022-07-22 中国联合重型燃气轮机技术有限公司 透平动叶和燃气轮机
CN115163203A (zh) * 2022-06-24 2022-10-11 中国船舶重工集团公司第七0三研究所 一种具有空心叶冠结构的涡轮动叶
CN121803307A (zh) * 2026-03-10 2026-04-07 中国航发沈阳发动机研究所 一种高压涡轮导向叶片双层壁结构

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Publication number Publication date
EP1557533B1 (fr) 2008-03-12
DE502004006484D1 (de) 2008-04-24

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