EP2354453A1 - Turbinenmotorkomponente zur adaptiven Kühlung - Google Patents

Turbinenmotorkomponente zur adaptiven Kühlung Download PDF

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Publication number
EP2354453A1
EP2354453A1 EP10152428A EP10152428A EP2354453A1 EP 2354453 A1 EP2354453 A1 EP 2354453A1 EP 10152428 A EP10152428 A EP 10152428A EP 10152428 A EP10152428 A EP 10152428A EP 2354453 A1 EP2354453 A1 EP 2354453A1
Authority
EP
European Patent Office
Prior art keywords
depression
turbine engine
engine component
back face
body portion
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
EP10152428A
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English (en)
French (fr)
Other versions
EP2354453B1 (de
Inventor
Andrew Shepherd
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
Priority to EP10152428.8A priority Critical patent/EP2354453B1/de
Priority to US13/017,350 priority patent/US20110189015A1/en
Publication of EP2354453A1 publication Critical patent/EP2354453A1/de
Application granted granted Critical
Publication of EP2354453B1 publication Critical patent/EP2354453B1/de
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/18—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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/90—Coating; Surface treatment
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/232—Three-dimensional prismatic conical
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/292—Three-dimensional machined; miscellaneous tapered
    • 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
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/11—Purpose of the control system to prolong engine life

Definitions

  • the invention relates to adaptive cooling of a hot turbine engine component, particularly a turbine aerofoil within a turbine section of a gas turbine engine.
  • Components such as gas turbine blades, vanes and others in a turbine or a combustion section of a gas turbine may be affected by hot fluids resulting in oxidation of the components. This problem is addressed in several ways, e.g. by using heat resistant materials or by applying a coating, like a thermal barrier coating (TBC). Alternatively or additionally the components may be cooled. For this additional cooling, cooling air may be supplied and cooling features and cooling holes may be designed in regions to be cooled.
  • TBC thermal barrier coating
  • cooling may be necessary, the existence of cooling holes and the supply of cooling air to the region to be cooled - the cooling air may typically be taken from a compressor section of the gas turbine - may have negative side effects on the overall performance of the compressor and/or of the turbine section.
  • a gas turbine may be designed such that only that amount of cooling air is provided and distributed that is considered necessary.
  • material may still oxidise during operation, e.g. due to high temperature peaks.
  • an apparatus for adaptive cooling comprising a first component having at least one aperture extending therethrough with a sacrificial component positioned within the at least one aperture.
  • the first component is operable at a maximum duty temperature and the sacrificial component has a melting or sublimation point below the maximum duty temperature of the first component.
  • the sacrificial component defines an effective aperture, the size of which may be increased if, in use, the sacrificial component is subjected to a temperature between the melting or sublimation point of the sacrificial component and the maximum duty temperature of the first component.
  • the sacrificial component is a separate component than the first component.
  • Patent US 7,241,107 B2 shows that cooling air passages may be coated.
  • the coating is being made of a material that has an oxidizing property such that the material oxidizes away and opens the passage to more flow when exposed to a temperature above a critical temperature of an airfoil.
  • the airfoil surface is not properly cooled by a flow passing through the passage, the material oxidizes away until the size of the passage increases to allow for the proper amount of cooling air to flow to cool the airfoil.
  • the passages are through holes and each have a fixed diameter. To coat these passages may be difficult.
  • a leading edge of an airfoil may erode or corrode such that a passage for cooling air - the passage being built inside of the airfoil - may become exposed to allow a cooling fluid to effect film cooling of a selected portion of the airfoil.
  • the present invention seeks to find a solution in which the time of operation can be stretched further and downtime may be avoided even when oxidation of parts has already occurred, so that repair or replacement of parts can be performed at the next maintenance schedule and still providing a mode of operation with only marginal performance losses.
  • a turbine engine component particularly an aerofoil, having a body comprising a first surface exposed to a working fluid of high temperatures during operation, so that erosion and/or corrosion, e.g. particularly oxidation, may take place on the first surface.
  • the body is further comprising a second surface comprising at least one depression, the depression being exposed to a cooling fluid during operation, the depression being oriented such that, starting from the second surface, it deepens in direction of a back face of the first surface.
  • a depression may be in form of a blind hole or a trench.
  • the body also comprises a body portion between the back face and the first surface, wherein a diameter of the depression decreases from the second surface in direction of the back face.
  • the depression can be seen as to have a particularly V-shaped cross section.
  • the turbine engine component may be, besides the already mentioned aerofoils like blades or vanes in a turbine section of a gas turbine engine, a component inside a hot fluid path or a component defining and guiding the hot fluid path.
  • This component may be located in the turbine section, in a combustor section of a gas turbine engine or any intermediate component - like a transition duct between combustor and turbine section - or downstream component - like an exhaust or a diffuser of the gas turbine engine that is downstream of the turbine section.
  • the turbine engine component may be any component within a gas turbine engine or a similar type of engine having surfaces over which hot gases are swept. Furthermore these surfaces may be part of a vane aerofoil and/or a vane platform within a turbine section, and/or a blade aerofoil and/or a blade platform and/or a blade shroud within the turbine section. The surface may also be part of a combustor. Additionally the surface may be part of a static shroud and/or a heat shield. Besides, the invention is also applicable to other hot gas swept surfaces downstream of the turbine section in a diffuser and/or in an exhaust, including heat recovery system.
  • the depression may be casted or machined. If instead of a blind hole a through hole is produced during manufacturing, the through hole may be filled again partly, covered and/or plugged so that the first surface may be closed again.
  • the through hole may be filled by a heat resistant material, possibly by the same material as the body, or of a material with a melting point that will permanently exceed the temperature during operation.
  • a heat resistant material possibly by the same material as the body, or of a material with a melting point that will permanently exceed the temperature during operation.
  • this heat resistant material is not designed to melt during operation. But the surface directed to the hot fluid path of the heat resistant material may encounter oxidation.
  • the invention is not directed to oxidise surfaces of cooling passages. It is directed to scenarios in which a hot surface that is designed to guide the hot fluid - the first surface according to the invention - may oxidise.
  • the invention is advantageous, if the first surface oxidises until the body portion between the back face and the first surface gets oxidised away, so that the depression becomes a passage to the hot fluid path. Due to the fact that the depression is being exposed to the cooling fluid, this newly built passage performs the function of a cooling hole via which the cooling fluid will be guided in direction of the first surface. This results in cooling of the first surface and its surroundings in a proximity of the passage. As a consequence the oxidation may stop in that area.
  • the depression has a specific shape, so that the diameter of the depression decreases from the second surface in direction of the back face of the first surface.
  • this passage has an exit with a first diameter, as defined by the width of the depression, providing a first amount of cooling fluid. This will stop or slow down further oxidation in the future. Assuming, oxidation will not be prevented completely because the first amount of cooling fluid may not be sufficient to do so, the first surface including the rim of the newly built passage may oxidise further. This results in a further gradual reduction of surface material and also in enlarging the diameter of the cooling passage, due to the specific shape of the depression with its tilted walls.
  • a second amount of cooling fluid will be provided to the surroundings of the exit.
  • the second diameter will be greater than the first diameter, the second amount of cooling fluid will exceed the first amount of cooling fluid. This principle will continue until enough cooling fluid will pass the passage so that no oxidation will occur anymore in the surroundings of the passage.
  • the inner walls of the passage itself will not oxidise but the first surface will continue to oxidise which then directly has an effect on the diameter of the exit of the passage.
  • the invention advantageously provides a self adapting principle to adapt to the conditions within the hot fluid path.
  • the invention is advantageous because, if the heat will never reach an oxidising temperature, the depression will never become a cooling passage and therefore no performance losses will occur.
  • a part of the body adjacent to the body portion is of a further material that is less susceptible to erosion and/or corrosion, particularly oxidation, than the body portion.
  • the compositions of the material may be seamlessly change so that this effect may take place.
  • the part of the body adjacent to the body portion may be coated.
  • the last paragraph is based on the fact that one might anticipate at which positions hot spots will occur. This may not always be the case.
  • the invention is also advantageous if the positions of hot spots will not be known beforehand or on which components the hot spots will occur. This may be the case if the locations of the hot spots depend on how the parts are assembled. In this case a plurality of depressions may be arranged in many places on several components and only these depressions will become holes eventually that are actually located at or near a hot spot.
  • the body portion may also incorporate a protective coating over the first surface. Cooling holes would then be exposed after the coating and part of the component material have been lost.
  • FIG. 1 one half of an aerofoil 1 of a turbine is shown in a perspective view as seen from the inside of the aerofoil 1.
  • the second half of the aerofoil 1 that would disallow the view to the inside of the aerofoil 1 is not shown is only indicated by broken and dashed lines.
  • the aerofoil 1 for example a guide vane or a blade within a turbine section of a gas turbine engine, represents the turbine engine component as defined in the invention and will be present in a working fluid flow 30 during operation.
  • This working fluid flow 30 is indicated as a double arrow and may be a hot gas provided by a combustion chamber located upstream.
  • the aerofoil 1 has a circumferential wall 2 as a body of turbine engine component and a cavity to allow cooling fluid - a cooling fluid flow 31 is indicated via a further arrow - to cool the wall 2 from the inside and to supply cooling holes (not shown in the figure).
  • the wall 2 has a first surface 10 which is exposed, in use, to the working fluid flow 30 - a hot gas path - of high temperatures.
  • the first surface 10 is directed to the outside of the aerofoil 1.
  • the wall 2 may be coated from the outside (not shown in Fig. 1 ; indicated in Fig. 2C ), for example by a thermal barrier coating or oxidation resistant coating, such as MCrAlY.
  • the first surface 10 may also comprise the coating.
  • the wall 2 has a second surface 11 which is facing away from the hot gas path and is being exposed to the cooling fluid flow 31.
  • the second surface 11 is the inner wall of the aerofoil 1, surrounding the cavity.
  • the second surface 11 comprises a plurality of depressions 20, 21 in different forms, e.g. a first depression 20 in form of a truncated cone, having a circular rim with the second surface 11.
  • a first depression 20 in form of a truncated cone, having a circular rim with the second surface 11.
  • a second depression 21 a trench is indicated having two side planes and a third bottom plane. Further types of depressions will be shown in Fig. 2 .
  • the depressions 20, 21 are being exposed to the cooling fluid during operation. At the time of installation, the depressions 20, 21 are blind holes and may not supply any cooling fluid to the first surface 10.
  • each of the depressions 20, 21 is oriented such that, starting from the second surface 11, the depression 20, 21 deepens in direction of a back face 12 of the first surface 10, and at the same time, a diameter of the depression 20, 21 decreases from the second surface in direction of the back face, thus resulting in a depression in form of a truncated cone (like the first depression 20), a truncated pyramid, a truncated triangular prism (like the second depression 21), or any kind of frustum, for example.
  • the back face 12 will be the result of the truncation.
  • the back face 12 may be substantially flat.
  • a diameter may be defined as the distance between two opposing points on a wall of the depression 20, 21.
  • the diameter of the depression 20, 21 decreases from the second surface 11 in direction of the back face 12, with a first diameter d1 at the second surface 11 and a second diameter d2 at the back face 12, with d1 greater than d2.
  • walls of the depression 20, 21 are tilted, narrowing the diameter until the back face 12 is reached, which defines the narrowest plane with the depression 20, 21.
  • a body portion 40 can be defined between the back face 12 of the first surface 10 and the first surface 10, which is part of the wall 2.
  • the body portion 40 may comprise less material than other areas of the wall 2 and may define a narrow section of the wall 2. This is intentional, so that in case of oxidation of the first surface 11 of the wall 2, the wall 2 may be breached at this location and will result in a passage (see Fig. 2A and Fig. 2C , reference sign 50) for cooling air.
  • Fig. 1 the aerofoil 1 is shown as manufactured. All depression 20, 21 may be blind holes which have not become passages yet.
  • Fig. 2A the same is shown for a short section of the wall 2 as a cross section from a direction indicated by the line A-A in Fig. 1 . Furthermore Fig. 2A also shows the appearance of the wall section once oxidation of the first surface 10 has taken place.
  • a section of the wall 2 is shown, with its first surface 10 that is directed to the hot gas path.
  • Hot streaming gas is indicated again via an arrow for the working fluid flow 30.
  • the first depression 20 in form of a truncated cone and a second depression 21 in form of a truncated triangular prism is shown, which both appear in a cross section as truncated trapezoid or trapezium.
  • the first depression 20 has an opening of the first diameter d1 then narrows until the back face 12 is reached with a diameter of the second diameter d2.
  • the width of the material of the body portion 40 is indicated by the width w and may be a fraction of a width of a medium distance between the first surface 10 and the second surface 11.
  • the first drawing of Fig. 2A shows the aerofoil 1 without being oxidised, for example when manufactured or if installed in an area in which no extreme temperatures occur.
  • the second drawing of Fig. 2A show the aerofoil 1 after oxidisation has occurred for a longer time span.
  • the first surface 10 is affected by oxidisation resulting in material loss and in an oxidised first surface 13. If the temperature remains extremely hot, the oxidisation of the oxidised first surface 11 continues, until all the material of the body portion 40 has vanished. In this case the blind holes of the depressions 20, 21 will become through holes, i.e. passages 50 which allow cooling air to pass, as indicated via arrows for cooling fluid flow 31. As a further consequence the cooling air will cool the oxidised first surface 13 in a proximity of the passage 50.
  • the material degradation of the first surface 10 may continue at a slower pace. This result in further oxidisation, as indicated by a dashed line for a further oxidised first surface 14.
  • the height of the passages 50 will also decrease, which has a direct impact on the width of the passage 50, due to the tilted walls of the passage 50.
  • the throughput of cooling air will increase through the passage 50, because the second diameter d2 indicating the diameter of the passage exit will also increase. This has the consequence that the oxidisation of the first surface 10 will decrease further until a stable point of operation is reached, at which enough cooling air is provided via the passage 50.
  • the invention provides a self adjusting principle, in which cooling passages will be created “automatically” for hot spots and the turbine engine component will adapt the diameters of the exits of the cooling passages "itself” such that the minimum needed cooling air is provided, so that no oxidisation takes place any further.
  • Fig. 2B a number of depressions 22, 23, 24, 25 are shown, that are slight modifications of the first depression 20 and the second depression 21.
  • Fig. 2A showed depressions 20, 21 that show a plane symmetry in regards of a plane through the centre of the back face 12, the plane being perpendicular to the first surface 10 extending perpendicular to the plane of projection. Due to the symmetry tilted walls 51 of the depression 20, 21 have the same length in the projection of Fig. 2A . Depressions with different walls 51 are shown for the depressions 22 and 24 in Fig. 2B . This allows to direct the depression in a wanted direction, so that the cooling fluid, once the passage 50 has been occurred, will be guided in a the wanted direction, for example to provide film cooling or alternatively to create turbulences.
  • Depression 23 shows a specific form, in which the depression is a cone or a trapezoid without truncation.
  • the back face 12 may only be a point or a line. This will result in a very small passage 50 at the beginning.
  • depression 25 shows a different form, in which the walls 51 are convex.
  • Other forms may be also advantageous.
  • the wall 2 of the aerofoil 1 may be designed such that the oxidation mainly takes place or will be affecting the material faster on the first surface in the area of the body portion 40. This may be possible if a part of the wall 2 adjacent to the body portion 40 is of a further material that is less susceptible to erosion and/or corrosion, e.g. oxidation, than the body portion 40. This allows creating a passage 50 for cooling air even before the remainder of the first surface 10 of the wall 2 is totally oxidised.
  • Fig. 2C shows an embodiment in which the complete first surface 10 is coated.
  • the body portion 40 is sealing the passage 50 to become a depression in form of a blind hole merely by a coating 60.
  • the principle of operation does not differ from the previously said, thus resulting in a passage 50 as shown in the second drawing of Fig. 2C .
  • the solution of Fig. 2C may be advantageous in respect of manufacturing the aerofoil with its depressions, because the depression can be manufactured initially as through hole and then be closed by the coating 60.
  • the aerofoil with depressions may be manufactured by casting the aerofoil including its depressions or by casting a solid aerofoil without depressions and later drilling the depressions.
  • Fig. 2D shows a slightly different embodiment, in which a depression 27 is a blind hole as in Fig. 2A , with an additional coating 60 as known from Fig. 2C .
  • the body portion 40 which may oxidise away, comprises a portion of the coating 60 and a portion of the wall 2 between the back face 12 and the first surface 10, as indicated in Fig. 2D .
  • the part of the aerofoil 1 is shown once at the beginning of operation and in a second drawing after oxidisation has taken place and the passage 50 has been formed.
  • This invention is particularly advantageous, because the operation time of turbine engine components that may be affected by oxidisation may be extended. Furthermore the invention is directed to protect installed components in a way that a slight oxidation of surfaces will be accepted. Besides, an adaption to the needed amount of cooling fluid is possible, and due to the fact that the walls of the depression are tilted, this adaption may take less time than in other implementations.
  • the invention may have further advantages in respect of internal convection cooling.
  • One of the benefits of a tapered depression is that it increases the internal surface area of the wall to which it is applied. This means that the depression provides some cooling benefit even before the outside part - the body portion 40 - of the depression is removed and a passage is built. If there is active impingement cooling on the inside surface, as there often is in an aerofoil, this cooling benefit could be significant. In this case, to distinguish from prior art convective cooling solutions, it may be the target to design the depression such that in a perfect installation the convective cooling may be sufficient and no erosion and/or corrosion takes place.
  • the body portion 40 may be sized such that there is a chance that a passage 50 may arise during typical time of operation of the component.
  • the body portion 40 may have a lesser width than known depressions that are solely intended for convection cooling.
  • the convective heat transfer coefficient will also be enhanced due to turbulences of the cooling air resulting of the tapered depression.
  • the depression may be arranged such that the body portion may be eroded and/or corroded completely and generating the passage 50 before a next scheduled inspection of the gas turbine in cases of continuously occurring hot spots.
  • the invention allows easy manufacturing.
  • the turbine engine component may be easier to cast.
  • a ceramic forming the depression during the casting process is stronger with a tapered shape than with a parallel sided shape.
  • a depression could start from the second surface with a cylindrical shape, later followed by a tapered portion in direction of the back face.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP10152428.8A 2010-02-02 2010-02-02 Turbinenmotorkomponente zur adaptiven Kühlung Not-in-force EP2354453B1 (de)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP10152428.8A EP2354453B1 (de) 2010-02-02 2010-02-02 Turbinenmotorkomponente zur adaptiven Kühlung
US13/017,350 US20110189015A1 (en) 2010-02-02 2011-01-31 turbine engine component for adaptive cooling

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP10152428.8A EP2354453B1 (de) 2010-02-02 2010-02-02 Turbinenmotorkomponente zur adaptiven Kühlung

Publications (2)

Publication Number Publication Date
EP2354453A1 true EP2354453A1 (de) 2011-08-10
EP2354453B1 EP2354453B1 (de) 2018-03-28

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EP (1) EP2354453B1 (de)

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EP2873806A1 (de) * 2013-11-14 2015-05-20 Siemens Aktiengesellschaft Durch Wärmedämmschicht verbesserte Kühlanordnung für eine Turbomaschinenkomponente
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US10006367B2 (en) * 2013-03-15 2018-06-26 United Technologies Corporation Self-opening cooling passages for a gas turbine engine
US9394798B2 (en) * 2013-04-02 2016-07-19 Honeywell International Inc. Gas turbine engines with turbine airfoil cooling
JP6227653B2 (ja) * 2013-07-30 2017-11-08 三菱日立パワーシステムズ株式会社 蒸気タービンの水分除去装置、及びスリット孔の形成方法
US9784123B2 (en) 2014-01-10 2017-10-10 Genearl Electric Company Turbine components with bi-material adaptive cooling pathways
DE102014207790A1 (de) * 2014-04-25 2015-10-29 Siemens Aktiengesellschaft Kühlfluidkanal
US9718735B2 (en) * 2015-02-03 2017-08-01 General Electric Company CMC turbine components and methods of forming CMC turbine components
US10738622B2 (en) 2016-08-09 2020-08-11 General Electric Company Components having outer wall recesses for impingement cooling
US10508553B2 (en) 2016-12-02 2019-12-17 General Electric Company Components having separable outer wall plugs for modulated film cooling
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EP2857636A1 (de) * 2013-10-01 2015-04-08 Siemens Aktiengesellschaft Verbesserte Kühlanordnung für eine Turbomaschinenkomponente
EP2873806A1 (de) * 2013-11-14 2015-05-20 Siemens Aktiengesellschaft Durch Wärmedämmschicht verbesserte Kühlanordnung für eine Turbomaschinenkomponente
WO2015071141A1 (en) * 2013-11-14 2015-05-21 Siemens Aktiengesellschaft A thermal barrier coating enhanced cooling arrangement for a turbomachine component
EP2937512A1 (de) * 2014-04-23 2015-10-28 United Technologies Corporation Bauteil eines Gasturbinenkraftwerks und zugehörige Anordnung
US9797260B2 (en) 2014-04-23 2017-10-24 United Technologies Corporation Engine component with wear surface protection
US10662779B2 (en) 2016-11-17 2020-05-26 Raytheon Technologies Corporation Gas turbine engine component with degradation cooling scheme
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EP3409890A1 (de) * 2017-05-31 2018-12-05 General Electric Company Komponente zur verwendung im heissgaspfad einer industriemaschine
CN108979731A (zh) * 2017-05-31 2018-12-11 通用电气公司 自适应地打开的冷却路径
CN108979733A (zh) * 2017-05-31 2018-12-11 通用电气公司 用于工业机器的热气路径中的部件
JP2019011753A (ja) * 2017-05-31 2019-01-24 ゼネラル・エレクトリック・カンパニイ 適応的に開口する冷却経路
EP3409894A1 (de) * 2017-05-31 2018-12-05 General Electric Company Adaptive abdeckung für einen kühlweg durch generative fertigung
US10704399B2 (en) 2017-05-31 2020-07-07 General Electric Company Adaptively opening cooling pathway
EP3409891A1 (de) * 2017-05-31 2018-12-05 General Electric Company Bauteile zur verwendung im heissgaspfad einer industriemaschine
US10927680B2 (en) 2017-05-31 2021-02-23 General Electric Company Adaptive cover for cooling pathway by additive manufacture
US11041389B2 (en) 2017-05-31 2021-06-22 General Electric Company Adaptive cover for cooling pathway by additive manufacture
WO2019074514A1 (en) * 2017-10-13 2019-04-18 General Electric Company COATED COMPONENTS HAVING ADAPTIVE COOLING OPENINGS AND METHODS OF MAKING SAME
US11352886B2 (en) 2017-10-13 2022-06-07 General Electric Company Coated components having adaptive cooling openings and methods of making the same

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