EP2857636A1 - Verbesserte Kühlanordnung für eine Turbomaschinenkomponente - Google Patents

Verbesserte Kühlanordnung für eine Turbomaschinenkomponente Download PDF

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Publication number
EP2857636A1
EP2857636A1 EP13186903.4A EP13186903A EP2857636A1 EP 2857636 A1 EP2857636 A1 EP 2857636A1 EP 13186903 A EP13186903 A EP 13186903A EP 2857636 A1 EP2857636 A1 EP 2857636A1
Authority
EP
European Patent Office
Prior art keywords
wall
shape
regions
recess
component according
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.)
Withdrawn
Application number
EP13186903.4A
Other languages
English (en)
French (fr)
Inventor
Katharina Bergander
Horst-Michael Dreher
Simon Maier
Khaled Maiz
Bettina Möller
Torsten Neddemeyer
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 EP13186903.4A priority Critical patent/EP2857636A1/de
Publication of EP2857636A1 publication Critical patent/EP2857636A1/de
Withdrawn legal-status Critical Current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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • 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

Definitions

  • the present invention relates to a blade or a vane for a turbomachine and more particularly to an arrangement for achieving continuous cooling of the blade or vane in case of a damage of a protecting coating of the blade or vane.
  • blade In the present application, for the sake of brevity only the term "blade” will be used, but the specifications and features can be transferred to a vane without further modifications. It might be mentioned that the basic idea of the invention is also applicable for a platform of a blade or a vane.
  • turbomachines various components of the turbomachine operate at very high temperatures.
  • turbine blades of the first stages of the turbomachine are thermally high loaded due to hot gas temperatures and high heat transfer coefficients caused by flow stagnation at the leading edges as well as high gas velocities.
  • the blades are coated with a ceramic thermal barrier coating (TBC) to reduce heat flux from the hot gas to the blade's base material which is typically a metal. This results in lower metal temperatures than without the TBC.
  • TBC ceramic thermal barrier coating
  • internal cooling is generally achieved by passing a cooling fluid through a core passage way cast into the blade component and into the airfoil of the blade, respectively.
  • the airfoil portion of the blade is cooled by directing the cooling fluid to flow through multiple flow paths of the core passage way that are designed to maintain all aspects of the turbine blade at a relatively uniform temperature.
  • the cooling fluid is directed over the internal surfaces of the airfoils to achieve a cooling effect.
  • film cooling holes are available through which the cooling fluid can exit the blade to create a separation in the form of a film of cold air between the hot gas flow and the blade surface. This results in external cooling of the blade.
  • the film cooling holes are especially located at thermally high loaded areas of the blade and the airfoil such as the leading edge. The cooling fluid can be fed to the film cooling holes by the core passage way.
  • the object is achieved by providing a component for a turbomachine according to claim 1.
  • the approach of the invention is to address the issue of local over temperature by implementing blind holes in predefined areas of certain regions of the component.
  • the certain regions can be located at an airfoil wall of a blade of the turbomachine.
  • the blind holes can be realized as predetermined breaking points of the blade walls by varying the thickness of the blade wall such that the wall comprises regions with less wall thickness.
  • the thinner walls of the blind holes will melt away faster than the material of the thicker walls surrounding the blind holes. This will open up the blind hole such that cooling fluid from the underlying core passage way can exit the airfoil, resulting in a film cooling effect.
  • This cooling will protect the blade at least temporarily and guarantee a proper functioning until the next inspection. This helps in preventing a damaged part from complete failure at least until the next inspection of the part.
  • the hole resulting from the molten material of the blind hole will act as a cooling opening, without affecting the structural integrity of the blade.
  • a component for a turbomachine comprises an outer wall having an inner surface and having a plurality of wall thicknesses, wherein the outer wall surrounds a volume.
  • the volume contains at least a section of a cooling arrangement, e.g. one of a plurality of cooling channels of the cooling arrangement which are used for guiding a cooling fluid.
  • the outer wall comprises first regions with at least one first wall thickness and second regions at defined locations of the outer wall with at least one second wall thickness. Each second region is surrounded by at least one of the first regions and the second wall thickness is less than the first wall thickness.
  • the outer wall in case the outer wall is exposed to the hot gas in the environment of the component during operation of the turbomachine due to a loss of TBC, the outer wall will melt first at the second region only to generate a hole in the outer wall.
  • the term "surrounded" includes an arrangement in which the second region is arranged in between two first regions, i.e. those first regions are arranged on either side of the particular second region. For example, this scenario might occur in case the second region comprises a lengthy slot, as will be described later, and the two first regions are located on both sides of the lengthy slot.
  • the inner surface of the outer wall is fluidly connected to the cooling path at least at the locations of the second regions, such that, in case the outer wall comprises the hole as described above at the location of the second region, the cooling fluid can stream through the hole to achieve a film cooling effect in the region surrounding the hole.
  • This film cooling effect protects the component against further melting such that a longer life time is achieved despite the loss of TBC.
  • At least one of the second regions comprises a recess with a three dimensional shape. Such a recess can be manufactured easily and it results in the reduced wall thickness in the second regions.
  • the three dimensional shape of the recess is such that a diameter of the recess increases with a decreasing thickness of the outer wall which might occur due to erosion and melting.
  • a reduced wall thickness results in a hole with a bigger diameter, allowing a stronger flow of cooling fluid through the hole. This measure achieves in an increased cooling effect.
  • the recess has a star-shaped cross section in a direction of viewing perpendicular to the outer wall. This increases the wetted surface in the recess and the hole, respectively, to that a stronger heat transfer can be achieved.
  • the three dimensional shape is a lengthy slot with an extension in a first direction along the outer wall being significantly larger than the extensions in the other directions.
  • the term "significantly" refers to dimensions of the extension in the first direction which are larger than the other two dimensions by a factor of 3 or more.
  • the extension in the first direction might be such that the lengthy slot forms a closed loop around the circumference of the component.
  • the three dimensional shape is a conic shape, a pyramidal shape, or a half dome shape.
  • the corresponding recesses are arranged such that the axis of symmetry of these shapes is perpendicular to the outer wall.
  • Such shapes can be manufactured easily and they achieve that the diameter of the recess increases with a decreasing thickness of the outer wall, resulting in a increased cooling effect with decreasing wall thickness.
  • the three dimensional shape is a truncated shape with a first flat surface A at a base of the shape and a second flat surface B at a top of the shape, wherein an area of the second surface B at the top of the shape is larger than an area of the first surface A at the base of the shape.
  • such shapes can be manufactured easily and they achieve that the diameter of the recess increases with a decreasing thickness of the outer wall, resulting in a increased cooling effect with decreasing wall thickness.
  • the three dimensional shape is cylinder or a box with at least two parallel surfaces A, B, wherein the three dimensional shape comprises a first flat surface A at a base of the shape and a second flat surface B at a top of the shape.
  • the corresponding recess would again be arranged such that the axis of symmetry of the cylinder is perpendicular to the outer wall at the location of the recess.
  • the recess is oriented such that the first surface A of the shape is facing to the inner surface of the outer wall and the second surface B is facing to an outer surface of the outer wall, wherein both the first surface A and the second surface B are essentially parallel to the outer wall.
  • the component typically comprises zones with different thermal loads, i.e. at least a zone with highest thermal load, a zone with medium thermal load, and a zone with lowest thermal load.
  • the second regions are located in the zone with highest thermal load. In that zone, the probability of damages is highest, so that the arrangement of the second regions in that zone guarantees the best protection with least efforts.
  • the component comprises a leading edge zone and a trailing edge zone, and the second regions are located in the leading edge zone.
  • the second wall thickness of one particular region of the second regions and the first wall thickness of the at least one first region surrounding the particular second region are dimensioned such that, upon unprotected exposure of the particular second region and the surrounding first region to a hot gas during operation of the turbomachine, the outer wall melts at the location of the particular second region to form a hole in the particular second region, such that a cooling fluid can leak out of the component to achieve the film cooing effect.
  • the wall thickness in that particular second region is such that the wall in that second region melts so much faster than the wall in the first region surrounding that only a hole is formed upon exposure of the outer wall to the hot gases. This would occur as soon as TBC is lost at the corresponding location. The above explained film cooling effect would then be achieved by cooling fluid leaking out of the hole.
  • the component can be an airfoil of a blade or a vane.
  • Each of the second regions may comprise a depression on the inner surface of the outer wall.
  • the outer surface of the outer wall can be even and the coating comprising a bond coat and the TBC can be applied easily.
  • a preferred method for manufacturing a component for a turbomachine according to the invention comprises
  • Embodiments of the present invention described below relate to a blade component in a turbomachine.
  • the details of the embodiments described in the following can be transferred to a vane component without modifications, that is the terms "blade” or "vane” can be used in conjunction, since they both have the shape of an airfoil with an integrated cooling arrangement in the form of a core passage way comprising one or multiple flow paths through which a cooling fluid is directed.
  • the turbomachine may include a gas turbine, a steam turbine, a turbofan and the like.
  • the present invention relates to a component of a turbomachine, especially to a blade.
  • the blade is connected to a rotor of said turbomachine, wherein the rotor with the blade is rotatable around an axis of rotation.
  • any term describing a direction like "radial” or “axial” is with reference to the axis of rotation of the rotor, i.e. a radial direction means a direction perpendicular to the axis of rotation of the rotor and an axial direction is in parallel to the axis of rotation.
  • FIG 1 shows a schematic diagram of an exemplary blade 1 of a rotor (not shown) of a turbomachine, such as a gas turbine.
  • the blade 1 includes an airfoil portion 20, a root portion 30, and a platform portion 40.
  • the airfoil portion 20 projects from the root portion 30 in a radial direction and the platform portion 40 is located between the airfoil portion 20 and the root portion 30.
  • the airfoil portion 20 extends radially along a longitudinal direction of the blade 1.
  • the blade 1 is attached to a body of the rotor (not shown), in such a way that the root portion 30 is attached to the body of the rotor whereas the airfoil portion 20 is located at a radially outermost position.
  • the platform portion 40 is attached to the radial outer surface of the rotor. Platform portions of neighbouring blades form an essentially cylindric surface.
  • the blade 1 and the airfoil portion 20 comprises a cooling arrangement 50 (not visible in FIG 1 ), typically with a plurality of cooling paths and cooling cavities.
  • a cooling fluid 59 is directed through the cooling arrangement 50 to maintain a suitable temperature of the blade 1 and the airfoil 20, respectively.
  • FIG 2 shows a cross-sectional view of the airfoil 20 in a radial direction, including a simplified version of the cooling arrangement 50 with a plurality of channels 51-56.
  • the airfoil portion 20 has a pressure side 22 and a suction side 23.
  • the pressure side 22 and the suction side 23 are joined together along a leading edge 24 and a trailing edge 25.
  • the leading edge 24 and the trailing edge 25 extend along the radial direction.
  • the airfoil 20 comprises an outer wall 21 with at least one inner surface 29.
  • the outer wall 21 surrounds a volume in which the cooling arrangement 50 is arranged.
  • the volume contains ribs 28, which are arranged to divide the cooling channels 51-56 inside the airfoil 20, wherein the ribs 28 are usually only slightly thicker than the outer wall 21.
  • the channels 51-56 of the cooling arrangement 50 might be interconnected in a serpentine manner or they are connected to a separating cavity via which the cooling fluid 59 would be provided.
  • the cooling arrangement 50 itself will not be described in more detail since its design is not an essential part of the invention. It is sufficient to mention that the cooling arrangement 50 has a cooling path along which the cooling fluid 59 is directed. The cooling path extends along the channels 51.
  • the cooling path and the channels 51 is fluidly connected with an inner surface 29 of the outer wall 21 of the airfoil 20, such that during operation of the turbomachine, when the cooling fluid 59 is streaming through the cooling arrangement 50 and through the channels 51 along the cooling path, the cooling fluid 59 is in connection with the inner surface 29 of the outer wall 21, so that a heat transfer from the inner surface 29 to the cooling fluid 59 is achieved.
  • one or more cooling channels might be arranged such that they are not fluidly connected to the inner surface 29 of the outer wall 21.
  • all channels 51 of the cooling arrangement 50 are fluidly connected to a section of the inner surface 29 of the outer wall 21.
  • the outer wall 21 comprises first regions 71 in which the outer wall 21 has a first wall thickness d1. Moreover, the outer wall 21 comprises second regions 72 at defined locations on the outer wall 21. In the second regions 72 the outer wall 21 has a wall thickness d2 which is less than the first wall thickness d1, i.e. d2 ⁇ d1. Thus, the outer wall 21 comprises a discontinuous wall thickness. In both cases, the wall thicknesses d1, d2 are measures for the extension of the outer wall 21 in a direction perpendicular to the outer wall 21.
  • the outer wall 21 may comprise recesses 73 of a certain three dimensional (3D) shape in the second regions 72.
  • the thickness d2 of the outer wall 21 is less than the wall thickness d1 in regions surrounding the recess 73.
  • the region surrounding the recess 73 in the second region 72 is the first region 71.
  • the inner surface 29 of the outer wall 21 is fluidly connected with the cooling arrangement 50 and the cooling channels 51-56, respectively.
  • the cooling fluid 59 gets in contact with the second regions 72.
  • the second regions 72 with the recesses 73 are located in a zone of the airfoil 20 which has the highest thermal load during operation of the turbomachine. Such a zone would be located at the leading edge 24.
  • the second regions 72 with reduced thickness d2 are preferably located at least at the inner surface 29 of the outer wall 21 at the particular cooling channel 51 which is located at the leading edge 24.
  • FIG 2 also shows that the outer wall 21 of the airfoil 20 is covered by a coating 60 which can comprise a bond coat 61 and a thermal barrier coating (TBC) 62.
  • a coating 60 which can comprise a bond coat 61 and a thermal barrier coating (TBC) 62.
  • TBC thermal barrier coating
  • the bond coat 61 is applied on the outer surface 27 of the outer wall 21 and the TBC 62 is applied on the bond coat 61, i.e. the bond coat 61 is arranged between the outer wall 21 and the TBC layer 62.
  • the material of the outer wall 21 may comprise one or more metals.
  • the material of the outer wall 21 would melt if it is exposed to the hot gases in the environment of the blade during operation of the turbomachine. Therein, the melting time necessary until a hole in the wall occurs depends on the thickness of the unprotected outer wall 21 at the location of exposure to the hot gas. Such an exposure would occur when the coating 60, especially the TBC 61, is damaged.
  • FIG 3 shows an rotated, enlarged view on the section III marked in FIG 2 with the outer wall 21, the coating 60 with bond coat 61 and TBC 62, hot gas 80, first and second regions 71, 72, recesses 73, and with the particular cooling channel 51.
  • FIG 4 depicts the situation in which the coating 60 is damaged and a part of the coating 60 has been lost.
  • both a part of the bond coat 61 and of the TBC 62 have been lost.
  • the material of the outer wall 21 at the location of the damage is exposed to hot gas 80 in the environment of the blade.
  • the material of the outer wall 21 at the location of the damage will start melting.
  • the wall thickness d2 in the second region 72 at the location of the recess 73 is less that the wall thickness d1 in the first region 71, i.e. in the region surrounding the recess 73, the wall material in the second region 72 will be molten away faster than the material in the surrounding region 71.
  • the damage would be detected with the next inspection interval and the blade would be repaired. However, it can be assured that the blade can be used for normal operation in spite of the local damage.
  • the airfoil 20 typically comprises zones with different thermal loads during operation of the turbomachine.
  • the thermal load will be highest in a zone around the leading edge 24 and lowest in a zone around the trailing edge 25.
  • thermal load will be medium.
  • the second regions 72 are located only in the zone with highest thermal load, i.e. in the zone around the leading edge 24. Additionally, second regions 72 might be located in the intermediate zone.
  • the distance between neighbouring second regions 72 should be less than 10mm in all directions along the outer wall 21. This is especially applicable in the leading edge zone.
  • the recesses 73 have a certain three dimensional (3D) shape.
  • a recess 73 can be cylindric, wherein the recess 73 would be arranged such that the axis of symmetry of the cylinder is perpendicular to the outer wall 21 at the location of the recess.
  • the recess 73 can be conical.
  • the conical shape is only a section of a full cone, i.e. a truncated cone or a conic section, as shown FIG 6 .
  • the shape has a first flat surface A at the base and a second flat surface B at the top.
  • the cross-sections of those surfaces A, B can be round or oval.
  • the area of the surface B at the top is less than the area of the surface A at the base of the conic section.
  • the recess 73 has a pyramidal shape.
  • the pyramidal shape is only a section of a full pyramid, i.e. a truncated pyramid or a pyramidal section, as shown FIG 7 .
  • the shape has a first flat surface A at the base and a second flat surface B at the top.
  • the cross-sections of those surfaces A, B can be rectangular, especially square.
  • the area of the surface B at the top is less than the area of the surface A at the base of the pyramidal section.
  • the recess 73 is box shaped.
  • the box might be cubic, cuboid, or rectangular cuboid, as shown FIG 8 .
  • the extension of the boxed shaped recess 73 in one particular direction parallel to the outer wall 21, i.e. to the inner surface 29 or the outer surface 27 of the outer wall 21, is substantially larger than the extensions in the other two directions, as shown in FIG 9 .
  • the recess 73 has the shape of a lengthy slot.
  • the extension in the particular direction might be such that the lengthy slot forms a closed loop around the circumference of the airfoil 20.
  • the recess 73 has the shape of a half dome.
  • the domed shape is only a section of a full half dome, i.e. a truncated half dome, as shown FIG 10 .
  • the shape has a first flat surface A at the base and a second flat surface B at the top.
  • the cross-sections of those surfaces A, B can be round or oval.
  • the area of the surface B at the top is less than the area of the surface A at the base of the conic section.
  • the recess is oriented such that the larger base surface A of the shape is facing to the inner surface 29 of the outer wall 21 and the smaller top surface B is facing at the outer surface 27 of the outer wall 21. Both the top surface B and the base surface A are essentially parallel to the outer wall 21.
  • the shape of the recess increases in diameter with a decreasing thickness of the outer wall 21 due to erosion and melting to increase cooling flow and thereby increase cooling.
  • the equivalent diameter of the cross-section area of the recess 73 can be between 0.0 and 0.7mm at the top and between 0.2 and 1.5mm at the base.
  • the whole cross-section of the recess 73 in a direction of viewing perpendicular to the outer wall 21 at the location of the recess 73 can be shaped like a star instead of a circle or a rectangle etc., as shown in FIG 11 in a direction of viewing perpendicular to the outer wall 21.
  • This measure increases the so called wetted surface, i.e. the surface of the film on the surface of the outer wall 21 at the location of the hole 74 and its surroundings, and, therewith, the heat transfer to the cooling fluid 59.
  • the first regions 71 are interconnected with each other, i.e. practically the first regions 71 form an uniform, extended surface and the second regions 72 with the recesses 73 are depressions in the uniform, extended surface 71.
  • the recesses 73 and the second regions 72, respectively, are not interconnected.
  • the second regions 72 can be distributed in the extended surface 71 according to a certain pattern.
  • the pattern can be a hexagonal pattern with the second regions 72 and the recesses 73 located on the corners of the hexagons of the pattern, as shown in FIG 12 .
  • the pattern might be a triangular ( FIG 13 ) or a quadratic pattern ( FIG 14 ) consisting of a plurality of regular triangles or squares, respectively, with the second regions 72 located at the corners of the triangles or squares.
  • FIGs 12, 13 , and 14 only few of the second regions 72 have been marked with reference signs.
  • the recesses 73 are located on the inner surface 29 of the outer wall 21, as shown in FIG 3 and 4 .
  • the outer surface 27 of the outer wall 21 is even.
  • the recesses 73 can be located on the outer surface 27 of the outer wall 21, as shown in FIG 15 .
  • the inner surface 29 of the outer wall 21 is even.
  • recesses 73 are located both on the outer surface 27 of the outer wall 21 and on the inner surface 29 of the outer wall 21, i.e. a recess 73 would have the form of a double-recess.
  • a recess 73-1 on the outer surface 27, i.e. an outer recess 73-1 opposes a recess on the inner surface, i.e. an inner recess 73-2, in a direction perpendicular to the outer wall 21 at the location of the recess.
  • the recesses 73-1, 73-2 are located such that a pair of recesses comprising an outer recess 73-1 and a corresponding inner recess 73-2 is located in one second region 72.
  • the thickness d2 of the outer wall 21 at the location of the double-recess 73 is less than the first wall thickness d1 in the first regions 71 of the outer wall 21.
  • the surface 27, 29 of the outer wall 21 comprising the recesses 73 might be compared with the surface of a golf ball.
  • the airfoil 20 with the outer wall 21 and the discontinuities 73 can be manufactured by either conventional casting using a specially modified mold or by creating the part not by casting but by a selective laser melting (SLM) process.
  • SLM selective laser melting
  • the mold can be manufactured either in a conventional way in one or more steps and pieces or the SLM process can be used to add certain features to the mold surface.
  • the formulation "perpendicular to the outer wall 21" is to be understood as “perpendicular to the outer surface 27 of the outer wall 21". Since the inner surface 29 and the outer surface 27 of the outer wall 21 are essentially parallel, the formulation can also be understood as “perpendicular to the inner surface 29 of the outer wall 21".

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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)
EP13186903.4A 2013-10-01 2013-10-01 Verbesserte Kühlanordnung für eine Turbomaschinenkomponente Withdrawn EP2857636A1 (de)

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EP13186903.4A EP2857636A1 (de) 2013-10-01 2013-10-01 Verbesserte Kühlanordnung für eine Turbomaschinenkomponente

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3409891A1 (de) * 2017-05-31 2018-12-05 General Electric Company Bauteile zur verwendung im heissgaspfad einer industriemaschine
CN108979726A (zh) * 2017-05-31 2018-12-11 通用电气公司 通过增材制造的用于冷却通路的自适应盖
JP2019015285A (ja) * 2017-05-31 2019-01-31 ゼネラル・エレクトリック・カンパニイ 付加製造による冷却経路用の適応カバー

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EP1669545A1 (de) * 2004-12-08 2006-06-14 Siemens Aktiengesellschaft Schichtsystem, Verwendung und Verfahren zur Herstellung eines Schichtsystems
US20090074576A1 (en) * 2006-04-20 2009-03-19 Florida Turbine Technologies, Inc. Turbine blade with cooling breakout passages
EP2354453A1 (de) * 2010-02-02 2011-08-10 Siemens Aktiengesellschaft Turbinenmotorkomponente zur adaptiven Kühlung
US20120156054A1 (en) * 2010-12-15 2012-06-21 General Electric Company Turbine component with near-surface cooling passage and process therefor

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1375825A1 (de) * 2002-06-17 2004-01-02 General Electric Company Ausfallsichere filmgekühlte Wand
EP1669545A1 (de) * 2004-12-08 2006-06-14 Siemens Aktiengesellschaft Schichtsystem, Verwendung und Verfahren zur Herstellung eines Schichtsystems
US20090074576A1 (en) * 2006-04-20 2009-03-19 Florida Turbine Technologies, Inc. Turbine blade with cooling breakout passages
EP2354453A1 (de) * 2010-02-02 2011-08-10 Siemens Aktiengesellschaft Turbinenmotorkomponente zur adaptiven Kühlung
US20120156054A1 (en) * 2010-12-15 2012-06-21 General Electric Company Turbine component with near-surface cooling passage and process therefor

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3409891A1 (de) * 2017-05-31 2018-12-05 General Electric Company Bauteile zur verwendung im heissgaspfad einer industriemaschine
CN108979726A (zh) * 2017-05-31 2018-12-11 通用电气公司 通过增材制造的用于冷却通路的自适应盖
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
JP2019015285A (ja) * 2017-05-31 2019-01-31 ゼネラル・エレクトリック・カンパニイ 付加製造による冷却経路用の適応カバー
JP2019023456A (ja) * 2017-05-31 2019-02-14 ゼネラル・エレクトリック・カンパニイ 付加製造による冷却通路のための適応カバー
US10704399B2 (en) 2017-05-31 2020-07-07 General Electric Company Adaptively opening cooling pathway

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