US20100068069A1 - Turbine Blade - Google Patents

Turbine Blade Download PDF

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Publication number
US20100068069A1
US20100068069A1 US12/447,972 US44797207A US2010068069A1 US 20100068069 A1 US20100068069 A1 US 20100068069A1 US 44797207 A US44797207 A US 44797207A US 2010068069 A1 US2010068069 A1 US 2010068069A1
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US
United States
Prior art keywords
heat shield
shield element
blade
turbine
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.)
Abandoned
Application number
US12/447,972
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English (en)
Inventor
Fathi Ahmad
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
Original Assignee
Siemens AG
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 filed Critical Siemens AG
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: AHMAD, FATHI
Publication of US20100068069A1 publication Critical patent/US20100068069A1/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
    • 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

Definitions

  • the invention refers to a turbine blade according to the claims.
  • Turbomachines especially gas turbines, are used in many areas for driving generators or driven machines.
  • a gas turbine customarily has a rotatably mounted rotor which is enclosed by a fixed casing.
  • the fixed sub-assemblies of a gas turbine are also collectively referred to as a stator.
  • the energy content of a fuel in this case is used for producing a rotational movement of the rotor components.
  • the fuel is combusted in a combustion chamber, wherein compressed air is supplied by an air compressor.
  • the operating medium which is produced in the combustion chamber as a result of the combustion of the fuel, being under high pressure and at high temperature, is guided in the process through a turbine unit which is connected downstream to the combustion chamber, where it is expanded, performing work.
  • a number of rotor blades which are customarily assembled in blade groups or blade rows, are arranged on these rotor components and drive the rotor components by means of impulse transmission of the flow medium.
  • stator blade rows which are connected to the turbine casing, are moreover customarily arranged between adjacent rotor blade rows.
  • the turbine blades, especially the stator blades in this case customarily have a blade airfoil, which is extended along a blade axis, for suitable guiding of the operating medium and upon which a platform, which extends transversely to the blade axis, can be formed on the end face for fastening the turbine blade on the respective carrier body.
  • the components and component parts of the gas turbine which are exposed to these high temperatures of the operating medium are therefore subjected to a high thermal stress.
  • the affected components especially the rotor blades and/or stator blades of the turbine unit, are cooled.
  • the turbine blades in this case are customarily provided with cooling passages, wherein an effective and reliable cooling of the leading edge of the respective turbine blade, which is thermally stressed to a particularly high degree, is especially to be ensured.
  • cooling air is customarily used in this case.
  • This cooling air can be fed to the respective turbine blade via a number of cooling medium passages which are integrated into the blade airfoil or the blade profile. From these cooling medium passages the cooling air flows into discharge passages, which branch off from these, of the respectively provided regions of the turbine blades, as a result of which a convective cooling of the blade interior and of the blade wall is achieved.
  • these passages are left open so that after flowing through the turbine blade the cooling air flows from discharge openings, which are also referred to as film cooling holes, and form a cooling film on the surface of the blade airfoil.
  • this cooling air film By means of this cooling air film, the blade basic body is largely protected on the surface against a direct and intensive contact with the hot operating medium which flows past at high velocity.
  • the discharge openings in this region are customarily arranged uniformly along at least two rows which are oriented parallel to the leading edge.
  • the discharge passages, moreover, as a rule are oriented at an angle to the longitudinal direction of the turbine blade, which assists the forming of the protective cooling air film which flows along the surface.
  • the leading edge of the turbine blade can moreover be provided with a heat shielding coating.
  • This heat shielding coating expediently consists of a material which is more resistant to temperature than that of the blade basic body.
  • the heat shielding coating is characterized by a low coefficient of thermal conductivity, as a result of which the temperature stress of the base material of the blade body is reduced. Therefore, the service life of the turbine blade is increased as a result of such a heat shielding coating in conjunction with cooling of the leading edge region of the blade.
  • This heat shield has the disadvantage that after a certain time cracks occur in the heat shielding coating. These cracks reduce the protection of the blade basic body against the hot exhaust gas of the gas turbine so that as a consequence of the increased thermal stress crack development can also occur in the basic body of the turbine blade. Such cracks in the blade basic body endanger the operational safety and can lead to the breakdown of the gas turbine.
  • a modular turbine blade of the type referred to in the introduction is known from GB 841 117.
  • the turbine blade comprises a cast basic body with a blade airfoil upon which a plurality of cooling air blow-out slots are provided on the leading edge side, which are covered at a distance by a guard plate which is fastened on the blade profile on the side.
  • the cooling air which issues from the slots cools the guarded leading edge in the manner of an impingement cooling, and subsequent to the impingement cooling is deflected by the plate in such a way that it can leave the modular turbine blade in the region of the pressure-side surface and suction-side surface.
  • the invention is therefore based on the object of disclosing a turbine blade of the aforementioned type which with simple means ensures an especially high operational safety of the gas turbine, even when used in high flow temperatures.
  • the invention in this case is based on the consideration that particularly with regard to the operational safety and the economical efficiency of a gas turbine the turbine blades should have a service life which is as long as possible as a result of a suitably selected heat shield. At the same time, the fact that particularly the leading edge of the turbine blade is thermally severely stressed should especially be taken into consideration. This leading edge should therefore especially be protected.
  • the heat shield element being attached at a distance on the blade basic body in the leading edge region, as a result of which a direct contact of the heat shield element with the blade basic body is avoided.
  • its outside surface in the leading edge region is provided with a number of secondary cooling passages, wherein these extend from the main cooling passage to the outside surface of the blade basic body.
  • These secondary cooling passages are arranged in a uniformly distributed manner behind the heat shield element for effective cooling in the leading edge region of the blade basic body. Therefore, stresses, and cracks which result from them, can be avoided.
  • this has a number of discharge passages which extend from its outside surface in the direction of the blade basic body.
  • This passage which is formed for guiding a cooling flow, additionally also serves as a connecting element between the heat shield element and the blade basic body.
  • the discharge passage in this case projects with one end into a main cooling passage which is formed inside the blade, wherein the medium which flows in the main cooling passage can flow for cooling the heat shield element on its outside surface.
  • the especially critical region that is to say the leading edge region of the turbine blade
  • the especially effective region is especially effectively protected against the high temperatures of the operating medium of the turbine.
  • the cooling is carried out in such a way by a cooling flow from the main cooling passage being directed in part through the discharge passages of the heat shield element onto its outside surface, and in part by a cooling flow from the main cooling passage flowing via the secondary cooling passages of the blade basic body through the gap which is formed by the heat shield element and the blade basic body.
  • the heat transfer between the heat shield element and the blade basic body is significantly reduced, moreover, by means of the cooling medium which flows between the inside surface of the heat shield element and the outside surface of the blade basic body by the heat in the leading edge region being carried away by means of the internal cooling flow.
  • the heat shield element especially preferably has a shape which is adapted to the profile of the blade basic body in the leading edge region. Consequently, the effect is achieved of the turbine blade also having a flow-optimized shape in the leading edge region after attaching the heat shield element. Moreover, the shape of the heat shield element which corresponds to the blade basic body leads to a uniform extent of the gap in the leading edge region. Consequently, the cooling medium flows with predominantly constant velocity along the outside surface of the blade basic body and along the inside surface of the heat shield element, as a result of which the cooling in the leading edge region of the turbine blade is carried out especially uniformly. Therefore, no excessively high stresses, which could lead to crack developments, occur, particularly in the blade basic body.
  • the heat shield element is produced from a material which is more resistant to temperature in comparison to the blade basic body. Since the heat shield element is directly exposed to inflow of hot operating medium during operation of the turbine, this component is particularly exposed to a high temperature stress. Therefore, the heat shield element should be produced from an especially temperature-resistant material in order to particularly ensure the operational safety and to minimize the downtimes of the turbine.
  • the heat shield element In addition to the use of temperature-resistant materials, for increasing the resistance of the heat shield element this should be cooled. If in this case the heat shield element is designed for impingement cooling for an especially effective cooling, this is achieved by the distance of the heat shield element from the blade basic body being sufficiently minimized.
  • the heat shield element is preferably arranged at a distance of 1 mm to 3 mm from the blade basic body.
  • a heat shield element which is attached up to this distance in the leading edge region of the turbine blade particularly ensures a sufficiently high impingement velocity of the cooling medium upon the inside surface of the heat shield element, as a result of which an especially effective cooling by means of impingement cooling is achieved. Since the static pressure in the main cooling passage of the blade basic body is predetermined, the impingement velocity of the cooling flow, in addition, for example, to the diameter of the secondary cooling passages, is determined in particular by the distance of the heat shield element from the blade basic body. A sufficiently high velocity of the cooling medium directly before impinging on the inside surface of the heat shield element is necessary since an intimate contact between the cooling medium and the inside surface of the heat shield element takes place in this way. By means of such impingement cooling a significantly more effective carrying away of heat is possible than is possible, for example, in the case of a film cooling.
  • the secondary cooling passages are arranged in a manner in which they are oriented essentially perpendicularly to the inside surface of the heat shield element. Therefore, the cooling flow from the main cooling passage impinges perpendicularly upon the inside surface of the heat shield element, as a result of which a large part of the kinetic energy of the cooling medium is used for an especially intimate contact between the particles in the cooling flow and the inside surface of the heat shield element. As a result, the heat of the heat shield element in an especially effective manner is transmitted to the internal cooling flow and carried away.
  • the heat shield element is connected to the blade basic body in the edge regions of the turbine blade.
  • the blade basic body preferably in its leading edge region, is provided with a recess.
  • the described heat shield element can advantageously be used at the places of the turbomachine where component parts and sub-assemblies of the thermal turbomachine are impinged upon with the hot operating medium.
  • the use of the heat shield element is especially preferable for the protection of the leading edge region of the turbine blade since the temperature stress of the blade basic body is especially high in this region.
  • the downtimes of the gas turbine are minimized by means of such a heat shield element since the service life is increased because of the heat shield element.
  • the advantages which are achieved with the invention are especially that by means of the heat shield element which is connected upstream to the blade basic body of the turbine blade an efficient protection of the leading edge region of the turbine blade against the high temperatures of the operating medium of the turbines is provided.
  • a heat shield system enables the use of impingement cooling, as a result of which the heat shield element can be cooled in an especially effective manner.
  • the possible occurrence of cracks, which extend from the outside surface of the heat shield element and spread into the blade basic body is prevented.
  • the heat shield elements according to the invention can subsequently be attached on the turbine blades in a simple manner and with relatively little cost.
  • FIG. 1 shows a half-section through a gas turbine
  • FIG. 2 shows a turbine blade, which is provided with a heat shield, in longitudinal section,
  • FIG. 3 shows a heat shield element which is sectioned in the longitudinal direction
  • FIG. 4 shows a heat shield element in cross section
  • FIG. 5 shows a cross section through a turbine blade which is provided with a heat shield element
  • FIG. 6 shows in alternative embodiment a turbine blade with a heat shield element which is integrated into the leading edge region of the blade basic body.
  • the gas turbine 1 has a compressor 2 for combustion air, a combustion chamber 4 , and also a turbine unit 6 for driving the compressor 2 and for driving a generator, which is not shown, or a driven machine.
  • the turbine unit 6 and the compressor 2 are arranged on a common turbine shaft 8 , which is also referred to as a turbine rotor, to which the generator or the driven machine is also connected, and which is rotatably mounted around its center axis 9 .
  • the combustion chamber 4 which is designed in the style of an annular combustion chamber, is equipped with a number of burners 10 for combustion of a liquid or gaseous fuel.
  • the turbine unit 6 has a number of rotatable rotor blades 12 which are connected to the turbine shaft 8 .
  • the rotor blades 12 are arranged on the turbine shaft 8 in the manner of a ring and so form a number of rotor blade rows.
  • the turbine unit 6 comprises a number of fixed stator blades 14 which are fastened also in the manner of a ring on an inner casing 16 of the turbine unit 6 , forming stator blade rows.
  • the rotor blades 12 in this case serve for driving the turbine shaft 8 by means of impulse transmission by the operating medium M which flows through the turbine unit 6 .
  • the stator blades 14 serve for flow-guiding of the operating medium M between two rotor blade rows or rotor blade rings which follow each other in each case as seen in the flow direction of the operating medium M.
  • a pair consisting of a ring of stator blades 14 or a stator blade row, and a ring of rotor blades 12 or a rotor blade row, which follow each other, in this case is also referred to as a turbine stage.
  • Each stator blade 14 has a platform 18 which is arranged as a wall element for fixing the respective stator blade on the inner casing 16 of the turbine unit 6 .
  • the platform 18 as also the turbine blade 12 , 14 , in this case is a comparatively thermally severely stressed component part.
  • Each rotor blade 12 is fastened in a similar manner on the turbine shaft 8 via a platform 19 which is also referred to as a blade root.
  • each guide ring 21 is arranged in each case on the inner casing 16 of the turbine unit 6 .
  • the outer surface of each guide ring 21 in this case is also exposed to the hot operating medium M which flows through the turbine unit 6 , and in the radial direction is at a distance from the outer end of the rotor blades 12 , which lie opposite it, by means of a gap.
  • the guide rings 21 which are arranged between adjacent stator blade rows in this case serve especially as shroud elements which protect the inner casing 16 , or other installed parts on the casing, against overstressing as a result of the operating medium M which flows through the turbine 6 .
  • the combustion chamber 4 in the exemplary embodiment is designed as a so-called annular combustion chamber, in which a multiplicity of burners 10 , which are arranged around the turbine shaft 8 in the circumferential direction, open into a common combustion space.
  • the combustion chamber 4 is designed in its entirety as an annular structure which is positioned around the turbine shaft 8 .
  • the combustion chamber 4 is designed for a comparatively high temperature of the operating medium M of about 1000° C. to 1600° C.
  • the rotor blades 12 as shown in FIG. 2 , have a heat shield element 22 which is attached in the leading edge region.
  • Each of the heat shield elements 22 which are attached to the rotor blades 12 is equipped on the operating medium side with an especially heat-resistant protective coating, such as ceramic, or is produced from a high temperature-resistant material.
  • the turbine blade 12 , 14 is provided with a number of secondary cooling passages 24 in the leading edge region.
  • the discharge passages 28 which are also attached in the leading edge region of the turbine blade 12 , 14 and which project into a main cooling passage 26 , serve as fastening elements for the heat shield element 22 in addition to the guiding of the cooling medium K.
  • the cooling air K which is preferably used as cooling medium K, flows via the secondary cooling passages 24 into the gap which is formed between the outside surface 32 of the blade basic body 30 and the inside surface 34 of the heat shield element 22 , and also through the discharge passages 28 of the heat shield element 22 , wherein the cooling air K which flows from the discharge passages 28 forms a protective film between the operating medium M and the outside surface 36 of the heat shield element 22 .
  • the cooling air K which escapes from the secondary cooling passages 24 of the blade basic body 30 flows against the inside surface 34 of the heat shield element 22 and cools this by means of the impingement cooling effect which occurs as a result.
  • FIGS. 3 and 4 show the heat shield element 22 in two different sectional views in each case, wherein it becomes apparent from the longitudinal section of the heat shield element 22 which is shown in FIG. 3 that the discharge passages 28 , as seen in the longitudinal direction of the heat shield element 22 , are arranged one behind the other, and wherein each discharge passage 28 extends from the outside surface 36 of the heat shield element 22 towards its inside surface 34 .
  • the discharge passages 28 in this case, as shown in FIG. 4 can be concentrically arranged perpendicularly to the longitudinal direction of the heat shield element 22 .
  • the heat shield element 22 has a shape which is adapted to the profile of the blade basic body 30 in the leading edge region. Consequently, the effect is achieved inter alia of the turbine blade 12 , 14 also having a flow-optimized shape after attaching the heat shield element 22 on the blade basic body 30 . Moreover, a heat shield element 22 which is curved in such a way results in a constant distance between the inside surface 34 of the heat shield element 22 and the outside surface 32 of the blade basic body 30 , as a result of which an especially effective cooling in this region is made possible.
  • the cooling air K which is required for the cooling in this case flows from the main cooling passage 26 of the turbine blade 12 , 14 , through the secondary cooling passages 24 , and through the discharge passages 28 , as a result of which a cooling film is formed on the outside surface 36 of the heat shield element 22 on account of the cooling air K which flows from the discharge passage 28 and on account of the operating medium M which flows in the turbine unit 6 .
  • the cooling of the inside surface 34 of the heat shield element 22 and of the outside surface 32 of the blade basic body 30 in the leading edge region of the turbine blade 12 , 14 is carried out by the discharging of the cooling air K from the secondary cooling passages 24 , wherein the inside surface 34 of the heat shield element 22 is cooled in an especially effective manner as a result of the impingement cooling effect which occurs in the process.
  • the secondary cooling passages 24 are preferably arranged in such a way that the cooling air K which flows from the secondary cooling passages 24 impinges perpendicularly to the inside surface 34 of the heat shield element 22 .
  • the distance of the heat shield element 22 from the blade basic body 30 in this case is preferably to be selected so that as a result of a sufficiently high flow velocity of the cooling medium K when impinging upon the inside surface 34 of the heat shield element 22 an intimate contact between the cooling air K and the impingement surface is brought about, and in this way the impingement effect is established.
  • FIG. 6 An especially expedient design of the turbine blade 12 , 14 with the heat shield element 22 according to the invention is shown in FIG. 6 .
  • the heat shield element 22 was integrated into the leading edge region of the blade basic body 30 , as a result of which the original external shape of the turbine blade 12 , 14 is advantageously maintained.
  • the aerodynamic design of the turbomachine is therefore not altered, as a result of which reduction of the efficiency of the gas turbine, for example on account of vortex formations on the outer edges when a heat shield element 22 is attached externally on the blade basic body 30 , is prevented.
  • the gap between the heat shield element 22 and the blade basic body 30 which is required for creating impingement cooling is consequently achieved in the case of this special embodiment of the turbine blade 12 , 14 by the heat shield element 22 being seated in a recess 38 which is provided in the blade basic body 30 .
  • the outside surface of the turbine blade 12 , 14 which reaches into the flow passage of the gas turbine is partially formed by the outside surface of the heat shield element 22 .
  • the free ends of the heat shield element 22 according to FIG. 5 are formed flush on the blade walls which are formed by the basic body 30 in the case of the design according to FIG. 6 in order to achieve an offset-free surface of the turbine blade 12 , 14 .
  • the part of the blade basic body 30 which lies opposite the heat shield element 22 is set back towards the inside of the blade so that the edge regions of the heat shield element 22 are connected to the blade body.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US12/447,972 2006-10-30 2007-09-20 Turbine Blade Abandoned US20100068069A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP06022622A EP1930544A1 (fr) 2006-10-30 2006-10-30 Aube de turbine
EP06022622.2 2006-10-30
PCT/EP2007/059989 WO2008052846A1 (fr) 2006-10-30 2007-09-20 Aube de turbine

Publications (1)

Publication Number Publication Date
US20100068069A1 true US20100068069A1 (en) 2010-03-18

Family

ID=37945313

Family Applications (1)

Application Number Title Priority Date Filing Date
US12/447,972 Abandoned US20100068069A1 (en) 2006-10-30 2007-09-20 Turbine Blade

Country Status (10)

Country Link
US (1) US20100068069A1 (fr)
EP (2) EP1930544A1 (fr)
JP (1) JP2010508461A (fr)
CN (1) CN101535601A (fr)
AT (1) ATE459784T1 (fr)
DE (1) DE502007003043D1 (fr)
ES (1) ES2341384T3 (fr)
PL (1) PL2084368T3 (fr)
RU (1) RU2405940C1 (fr)
WO (1) WO2008052846A1 (fr)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120160375A1 (en) * 2010-12-28 2012-06-28 Quinlan Yee Shuck Heat treating and brazing of an object
EP3105425A4 (fr) * 2014-02-13 2017-09-13 United Technologies Corporation Circuit de refroidissement pour composant de moteur à turbine à gaz, pourvu d'un élément de respiration en forme de piédestal
US20180135423A1 (en) * 2016-11-17 2018-05-17 General Electric Company Double impingement slot cap assembly
EP3431710A1 (fr) * 2017-07-19 2019-01-23 General Electric Company Écran de protection pour profil aérodynamique de moteur à turbine
US20200025382A1 (en) * 2017-09-29 2020-01-23 Doosan Heavy Industries & Construction Co., Ltd. Gas turbine
US20210332705A1 (en) * 2020-04-27 2021-10-28 Raytheon Technologies Corporation Airfoil with cmc liner and multi-piece monolithic ceramic shell

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CH700687A1 (de) * 2009-03-30 2010-09-30 Alstom Technology Ltd Gekühltes bauteil für eine gasturbine.
US10301945B2 (en) * 2015-12-18 2019-05-28 General Electric Company Interior cooling configurations in turbine rotor blades
CN105422188A (zh) * 2016-01-13 2016-03-23 北京航空航天大学 一种带隔热罩式复合冷却结构的涡轮叶片
CN115434756A (zh) * 2021-06-02 2022-12-06 中国航发商用航空发动机有限责任公司 涡轮叶片双层壁冷却结构

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US6241469B1 (en) * 1998-10-19 2001-06-05 Asea Brown Boveri Ag Turbine blade
US6709230B2 (en) * 2002-05-31 2004-03-23 Siemens Westinghouse Power Corporation Ceramic matrix composite gas turbine vane
US7871246B2 (en) * 2007-02-15 2011-01-18 Siemens Energy, Inc. Airfoil for a gas turbine

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US3836283A (en) * 1972-05-08 1974-09-17 Nat Aerospace Lab Construction of axial-flow turbine blades
US4026659A (en) * 1975-10-16 1977-05-31 Avco Corporation Cooled composite vanes for turbine nozzles
US6241469B1 (en) * 1998-10-19 2001-06-05 Asea Brown Boveri Ag Turbine blade
US6709230B2 (en) * 2002-05-31 2004-03-23 Siemens Westinghouse Power Corporation Ceramic matrix composite gas turbine vane
US7871246B2 (en) * 2007-02-15 2011-01-18 Siemens Energy, Inc. Airfoil for a gas turbine

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9399271B2 (en) * 2010-12-28 2016-07-26 Rolls-Royce Corporation Heat treating and brazing of an object
US20120160375A1 (en) * 2010-12-28 2012-06-28 Quinlan Yee Shuck Heat treating and brazing of an object
US10370981B2 (en) 2014-02-13 2019-08-06 United Technologies Corporation Gas turbine engine component cooling circuit with respirating pedestal
EP3105425A4 (fr) * 2014-02-13 2017-09-13 United Technologies Corporation Circuit de refroidissement pour composant de moteur à turbine à gaz, pourvu d'un élément de respiration en forme de piédestal
US10577942B2 (en) * 2016-11-17 2020-03-03 General Electric Company Double impingement slot cap assembly
CN108071426B (zh) * 2016-11-17 2022-07-05 通用电气公司 双重冲击槽盖组合件
WO2018093627A3 (fr) * 2016-11-17 2018-06-28 General Electric Company Ensemble capuchon à fente à double impact
EP3323988A1 (fr) * 2016-11-17 2018-05-23 General Electric Company Assemblage de refroidissement pour une aube de turbine
CN110168196A (zh) * 2016-11-17 2019-08-23 通用电气公司 用于涡轮翼型件的冷却组件以及涡轮组件的对应翼型件
CN108071426A (zh) * 2016-11-17 2018-05-25 通用电气公司 双重冲击槽盖组合件
US20180135423A1 (en) * 2016-11-17 2018-05-17 General Electric Company Double impingement slot cap assembly
EP3431710A1 (fr) * 2017-07-19 2019-01-23 General Electric Company Écran de protection pour profil aérodynamique de moteur à turbine
US20190024513A1 (en) * 2017-07-19 2019-01-24 General Electric Company Shield for a turbine engine airfoil
EP3650639A1 (fr) * 2017-07-19 2020-05-13 General Electric Company Écran de protection pour profil aérodynamique de moteur à turbine
US20200025382A1 (en) * 2017-09-29 2020-01-23 Doosan Heavy Industries & Construction Co., Ltd. Gas turbine
US11053850B2 (en) * 2017-09-29 2021-07-06 Doosan Heavy Industries & Construction Co., Ltd. Gas turbine
US20210332705A1 (en) * 2020-04-27 2021-10-28 Raytheon Technologies Corporation Airfoil with cmc liner and multi-piece monolithic ceramic shell
US11286783B2 (en) * 2020-04-27 2022-03-29 Raytheon Technologies Corporation Airfoil with CMC liner and multi-piece monolithic ceramic shell

Also Published As

Publication number Publication date
WO2008052846A1 (fr) 2008-05-08
DE502007003043D1 (de) 2010-04-15
PL2084368T3 (pl) 2010-07-30
EP2084368B1 (fr) 2010-03-03
RU2405940C1 (ru) 2010-12-10
ATE459784T1 (de) 2010-03-15
CN101535601A (zh) 2009-09-16
ES2341384T3 (es) 2010-06-18
EP1930544A1 (fr) 2008-06-11
EP2084368A1 (fr) 2009-08-05
JP2010508461A (ja) 2010-03-18

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