EP3450683A1 - Composant et procédé de fabrication associé - Google Patents

Composant et procédé de fabrication associé Download PDF

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Publication number
EP3450683A1
EP3450683A1 EP17189178.1A EP17189178A EP3450683A1 EP 3450683 A1 EP3450683 A1 EP 3450683A1 EP 17189178 A EP17189178 A EP 17189178A EP 3450683 A1 EP3450683 A1 EP 3450683A1
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EP
European Patent Office
Prior art keywords
section
component
ribs
surface roughness
turbo machine
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Withdrawn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
EP17189178.1A
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German (de)
English (en)
Inventor
Anthony Davis
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 EP17189178.1A priority Critical patent/EP3450683A1/fr
Publication of EP3450683A1 publication Critical patent/EP3450683A1/fr
Withdrawn legal-status Critical Current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
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/50—Intrinsic material properties or characteristics
    • F05D2300/516—Surface roughness

Definitions

  • the present disclosure relates to a component for a turbo machine, and a method of manufacturing a component for a turbo machine.
  • Gas turbines generally include a rotor with a number of rows of rotating rotor blades which are fixed to a rotor shaft and rows of stationary vanes between the rows of rotor blades which are fixed to the casing of the gas turbine.
  • a hot and pressurized working fluid flows through the rows of vanes and blades it transfers momentum to the rotor blades and thus imparts a rotary motion to the rotor while expanding and cooling.
  • the vanes are used to control the flow of the working medium so as to optimize momentum transfer to the rotor blades.
  • a typical gas turbine rotor blade comprises a root portion by which it is fixed to the rotor shaft, and an aerodynamically formed aerofoil portion which allows a transfer of momentum when the hot and pressurized working fluid flows along the aerofoil section.
  • Rotor blades tend to be hollow, for example comprising a plenum through which cooling air is forced.
  • the plenum may be divided by internal walls which are formed integrally with the aerofoil structure.
  • Typical cooling designs for nozzle guide vanes and turbine blades use turbulators (e.g. ribs, pins or pedestals) or impingement jets in various combinations to generate the necessary internal heat transfer.
  • a component (100) for a turbo machine comprising : a main body (104) having a fluid inlet (103) and fluid outlet (200); a cooling passage (204) extending between the fluid inlet (103) and the fluid outlet (200); the cooling passage (204) divided into a first section (204A) and a second section (204B) which extend between the fluid inlet (103) and fluid outlet (200); and at least part of the first section (204A) and/or second section (204B) comprise a surface roughness (Ra) no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • the surface roughness of the first section (204A) and second section (204B) may be different to one another.
  • the surface roughness of the remaining areas of the cooling passage (204) may be no less than about 1.5 ⁇ m but no more than about 3.5 ⁇ m.
  • the surface roughness of the first section (204A) and second section (204B) may be the same as one another.
  • the surface roughness may be defined by a plurality of spaced apart micro ribs (212) which extend at least part of the way across the cooling passage (204).
  • the micro ribs (212) may have a height and width of no less than 0.025mm and no greater than 0.1 mm.
  • micro ribs (212) may be polygonal in cross section.
  • One of the cooling passage (204) sections may be further provided with macro ribs (206) which extend across the cooling passage (204).
  • the macro ribs (206) may have a height and width of no less than 0.5mm and no greater than 5.0mm.
  • At least one micro rib (212) may be provided between adjacent macro ribs (206).
  • the macro ribs (206) may be polygonal in cross section.
  • the macro ribs (206) and micro ribs (212) may be are parallel with one another.
  • the macro ribs (206) and micro ribs (212) may be at an angle to one another.
  • the macro ribs (206) may be parallel to one another.
  • micro ribs (212) may be parallel to one another.
  • the component (100) may be one of : a rotor blade, stator vane or rotor disc.
  • a method of manufacturing a component (100) for a turbo machine comprising : providing a ceramic core element (300) for forming internal fluid flow passages of the component (100) by casting the component (100) around the ceramic core; wherein the walls of the ceramic core (300) which define a surface of the flow passages comprise a region having a predetermined surface roughness of no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • the surface roughness may be defined by micro grooves (312) provided in the surface of the ceramic core (300).
  • a component for a turbo machine for example a gas turbine engine, configured to have a cooling passage with a predetermined surface roughness which increases heat transfer between the material of the component and fluid/air passing through the component.
  • a method of making the component with the required pattern of surface roughness.
  • the present invention relates to a method of manufacture of a component for a turbo machine, and the component.
  • the turbo machine may be a gas turbine engine, and the component may be a rotor blade, stator vane or rotor disc.
  • Figure 1 shows an example of a gas turbine engine 60 in a sectional view, which illustrates the nature of components according to the present disclosure (for example rotor blades) and the environment in which they operate.
  • the gas turbine engine 60 comprises, in flow series, an inlet 62, a compressor section 64, a combustion section 66 and a turbine section 68, which are generally arranged in flow series and generally in the direction of a longitudinal or rotational axis 70.
  • the gas turbine engine 60 further comprises a shaft 72 which is rotatable about the rotational axis 70 and which extends longitudinally through the gas turbine engine 60.
  • the rotational axis 70 is normally the rotational axis of an associated gas turbine engine. Hence any reference to "axial”, “radial” and “circumferential" directions are with respect to the rotational axis 70.
  • the shaft 72 drivingly connects the turbine section 68 to the compressor section 64.
  • air 74 which is taken in through the air inlet 62 is compressed by the compressor section 64 and delivered to the combustion section or burner section 66.
  • the burner section 66 comprises a burner plenum 76, one or more combustion chambers 78 defined by a double wall can 80 and at least one burner 82 fixed to each combustion chamber 78.
  • the combustion chambers 78 and the burners 82 are located inside the burner plenum 76.
  • the compressed air passing through the compressor section 64 enters a diffuser 84 and is discharged from the diffuser 84 into the burner plenum 76 from where a portion of the air enters the burner 82 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 86 or working gas from the combustion is channelled via a transition duct 88 to the turbine section 68.
  • the turbine section 68 may comprise a number of blade carrying discs 90 or turbine wheels attached to the shaft 72.
  • the turbine section 68 comprises two discs 90 which each carry an annular array of turbine assemblies 12, which each comprises an aerofoil 14 embodied as a turbine blade 100.
  • Turbine cascades 92 are disposed between the turbine blades 100.
  • Each turbine cascade 92 carries an annular array of turbine assemblies 12, which each comprises an aerofoil 14 in the form of guiding vanes (i.e. stator vanes 96), which are fixed to a stator 94 of the gas turbine engine 60.
  • Figure 2 shows an enlarged view of a stator vane 96 and rotor blade 100.
  • Arrows “A” indicate the direction of flow of combustion gas 86 past the aerofoils 96,100.
  • Arrows “B” show air flow passages provided for sealing, and arrows “C” indicate cooling air flow paths for passing through the stator vanes 96.
  • Cooling flow passages 101 may be provided in the rotor disc 90 which extend radially outwards to feed an air flow passage 103 in the rotor blade 100.
  • the combustion gas 86 from the combustion chamber 78 enters the turbine section 58 and drives the turbine blades 100 which in turn rotate the shaft 72 to drive the compressor.
  • the guiding vanes 96 serve to optimise the angle of the combustion or working gas 86 on to the turbine blades.
  • Figure 3 shows a view of the rotor blades 100 looking upstream facing the flow "A" shown in Figure 2 .
  • Each rotor blade 100 comprises an aerofoil portion 104, a root portion 106 and a platform 108 from which the aerofoil extends.
  • the rotor blades 100 are fixed to the rotor disc 102 by means of their root portions 106, through which the flow passage 101 may extend.
  • the root portions 106 have a shape that corresponds to notches (or grooves) 109 in the rotor disc 90, and are configured to prevent the rotor blade 100 from detaching from the rotor disc 102 in a radial direction as the rotor disc 102 spins.
  • Figure 4 shows a part sectional view of a component according to the present disclosure.
  • the component is a rotor blade 100 as described above.
  • the component 100 comprises a main body 104, provided in this example as the aerofoil portion 104.
  • the main body 104 has a fluid inlet 103 which, as described above, in situ will be in flow communication with a cooling passage 101 or other fluid source.
  • the component may be another fluid/air cooled component for a gas turbine engine.
  • a nozzle guide vane or a turbine rotor in which the rotor blades are provided integrally with a rotor disc 90.
  • the inlet 103 may be provided as a single passage, or a plurality of passages.
  • the component 104 further comprises a fluid outlet, or a plurality of fluid outlets 200.
  • a cooling passage 204 extends between the fluid inlet 103 and fluid outlet 200.
  • the fluid outlet is provided along a trailing edge 202 of the aerofoil, for example as an elongate slit or plurality of openings.
  • the fluid outlet may also be through a hole in the blade tip section.
  • fluid inlet and fluid outlet may be taken to mean a single inlet and/or outlet, or a plurality of inlets and/or a plurality of outlets.
  • a subdivided inlet may feed the cooling passage 204 and/or a sub divided outlet may provide an exhaust path from the cooling passage 204.
  • the cooling passage 204 is divided into a first section 204A and a second section 204B.
  • the first section 204A and second section 204B extend between the fluid inlet 103 and fluid outlet 200.
  • the first section 204A and second section 204B are in series with each other between the fluid inlet 103 and fluid outlet 200.
  • the third section is in series with the first section and second section.
  • first section is intended to mean “different sections”.
  • third section is located downstream of the second section, and the second section is located downstream of the first section in the cooling passage, in terms of direction of flow of cooling flow.
  • first, second and third sections may be arranged differently.
  • the first section may be downstream of the second section, and/or the third section may be immediately downstream of the first section.
  • the first section 204A comprises turbulators 206 which extend from a wall of the rotor blade, providing a flow restriction and increased surface area in the flow path.
  • the second section 204B comprises a plurality of dividing walls 208 with spaces therebetween, which define flow passages and provide an increased surface area.
  • the third section 204C comprises pedestals 210 to provide an increased surface area.
  • Each of these macro cooling features are configured to increase surface area and promote turbulence and hence increase the amount of heat that will be transferred from the material of the rotor blade to the air passing therethrough.
  • cooling passage may not comprise any macro cooling features.
  • the surface roughness (Ra) of at least one of the first section and second section is configured to be no less than about 7 ⁇ m but no more than about 15 ⁇ m. That is to say, at least part of the first section and/or second section comprise a surface roughness (Ra) no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • the surface roughness (Ra) of at least one region of at least one section of the cooling passage 204 is configured to be no less than about 7 ⁇ m but no more than about 15 ⁇ m. That is to say at least one region of at least one section of the cooling passage 204 is configured to have a predetermined surface roughness (Ra) no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • the predetermined surface roughness (Ra) may be no less than about 8 ⁇ m but no more than about 11 ⁇ m.
  • the surface roughness of the remaining areas of the cooling passage 204 may be no less than about 1.5 ⁇ m but no more than about 3.5 ⁇ m.
  • the surface roughness of the second section may be no less than about 1.5 ⁇ m but no more than about 3.5 ⁇ m.
  • the surface roughness of the first section and second section may be the same as one another.
  • the first section is provided immediately downstream of the fluid inlet 103
  • the first section of the cooling passage may be located further downstream the cooling passage.
  • the predetermined surface roughness i.e. in the desired range
  • the predetermined surface roughness may be provided over the entire cooling passage 204 or to selected regions of the cooling passage 204.
  • the predetermined surface roughness i.e. in the desired range
  • the second section may be provided with the predetermined surface roughness in the desired range
  • the first section and third section may have a different surface roughness to the second section.
  • the surface roughness of the first section and second section may be different to one another.
  • the surface roughness is defined by a plurality of spaced apart micro ribs 212 which extend at least part of the way across the cooling passage 204, and in particular across the first section 204A of the cooling passage 204.
  • the micro ribs may have a height and width of no less than 0.025mm and no greater than 0.05mm.
  • the micro ribs may have a height and width of no less than 0.025mm and no greater than 0.1 mm. This is sufficient to create the surface roughness (Ra) of no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • the micro ribs 212 may be polygonal in cross section.
  • the micro ribs 212 may be square in cross section.
  • the macro ribs (turbulators) 206 may be polygonal in cross section, for example square.
  • the macro ribs may have a height and width of no less than 0.5mm and no greater than 5mm.
  • At least one micro rib may be provided between adjacent macro ribs.
  • the macro ribs 206 may be parallel with the micro ribs 212.
  • the macro ribs may be parallel to one another.
  • the micro ribs may be parallel to one another. At least one of the micro ribs may be angled relative to another one of the micro ribs.
  • the micro ribs may be provided 20° to 70° to the flow direction.
  • the macro ribs and micro ribs may be provided at an angle to one another. That is to say the macro ribs may be provided at an angle to the micro ribs.
  • FIGS. 6 to 10 show different arrangements according to the present disclosure. It will be appreciated that cooling passages through turbine blades may be provided in a great number of different ways, and that the provision of a predetermined surface roughness in the cooling passage may be applied regardless of the geometry of the cooling passage.
  • Figures 6 to 9 show variations on the arrangements shown in Figure 4 , with Figure 6 showing the first section 204A having the predetermined surface roughness (i.e. Ra provided as no less than about 7 ⁇ m but no more than about 15 ⁇ m).
  • the surface roughness of the remaining sections of the cooling passage have a different surface roughness to that of the first section 204A.
  • the surface roughness is provided by micro ribs which are perpendicular to the direction of flow through the first section 204A.
  • micro ribs are provided an angle to the direction of flow in the first section 204A.
  • the predetermined surface roughness may be provided in a region of the cooling passage 204 which extends through a region of the component which will, in use, require most cooling (for example, the leading edge region, as shown in Figures 4 to 10 ), with the surface roughness in the remainder of the cooling passage being provided with a surface roughness less than that having the predetermined surface roughness.
  • the majority of the cooling effect will be in the region having the predetermined surface roughness.
  • the component is a rotor blade, and hence the leading edge of the rotor blade will be the region requiring most cooling, it may be advantageous to provide the section of the cooling passage passing through the leading edge region with the predetermined surface roughness, as shown in Figures 4 to 10 .
  • the first section 204A and second section 204B of the cooling passage are provided with micro ribs which extend perpendicular to the flow of air through them.
  • the micro ribs may be provided at an angle to the direction of flow.
  • the surface roughness may be different in different sections or sub sections of the cooling passage 204.
  • the surface roughness may be configured to be no less than about 7 ⁇ m but no more than about 15 ⁇ m, and in another section the surface roughness has a different value (for example less than 7 ⁇ m or greater than 15 ⁇ m), and in the third section the surface roughness is less than in either of the other sections.
  • the surface roughness of two sections are provided with micro ribs to define the surface roughness
  • the third section 204C is not provided with micro ribs, although is provided with pedestals 210 as taught in the example of Figure 4 .
  • the predetermined surface roughness is different in all three of the sections 204A, 204B, 204C with a different pattern of micro ribs in each section.
  • the micro ribs are an angle to the direction flow
  • the micro ribs are provided in a crosshatch form (i.e.
  • micro ribs or surface finish may be provided in many ways to provide a predetermined surface roughness.
  • Figure 10 shows a further arrangement in which the cooling passage is divided into three sections, with a first section 204A being immediately downstream of a fluid inlet 103 and the flow being divided between a second section 204B and a third section 204C, where the flow passage through the second section 204B and third section 204C are arranged in parallel to the flow passage in the first section 204A.
  • the second section 204B is adjacent the trailing edge of the rotor blade
  • the third section 204C is adjacent the leading edge of the rotor blade.
  • the second section 204B has a fluid outlet 200 as well as being provided with micro ribs to define a predetermined surface roughness, for example no less than about 7 ⁇ m but no more than about 15 ⁇ m, with the micro ribs being provided at an angle to the longitudinal direction of the second section 204A.
  • the third section 204C has a first sub section 204C1 in flow communication with a second sub section 204C2 to form a "U" shape, the third section 204C being between the first section 204A and a leading edge of the rotor blade.
  • the second sub section 204C2 is provided with the predetermined surface roughness whereas the first sub section 204C1 is relatively smooth compared to the second sub section.
  • a region of predetermined surface roughness e.g. with a value of Ra in the range 7 ⁇ m to 15 ⁇ m, or in the range of in the range 8 ⁇ m to 11 ⁇ m
  • the rough surface can be used in isolation or in combination with conventional cooling methods to further enhance the heat transfer. That is to say, the predetermined surface roughness may be provided instead of macro cooling features, or in addition to macro cooling features.
  • surface roughness of cooling passages is provided in the range of 1.5 ⁇ m to no more than about 3.5 ⁇ m.
  • the component may be cast by casting the component around a ceramic core.
  • the method of manufacturing a component for a turbo machine may comprise the step of providing a ceramic core element 300, more than one element per rotor blade, for forming internal flow passages for example as shown in Figures 4 to 10 .
  • Figure 11 shows an example of a ceramic core 300 that may be used in the manufacture of a rotor blade according to the present disclosure.
  • Figure 12 shows an enlarged region of the features created in a cooling passage of the component by the features of the ceramic core 300 shown in Figure 11 .
  • the outer surfaces of the ceramic core 300 are provided with the predetermined surface roughness, which may be achieved in a variety of ways.
  • the walls of the ceramic core which define a surface of the flow passages may comprise a region or regions having a predetermined surface roughness of no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • the walls which define the ceramic core 300 which define a surface of the cooling passage sections may have a surface roughness of no less than about 7 ⁇ m but no more than about 15 ⁇ m.
  • a predetermined surface roughness may be provided on one or more of these sections 304A, 304B, 304C or a sub section thereof (for example a sub section of one of the sections, but not the whole section).
  • the surface roughness may be defined by micro grooves 312 provided in the surface of the ceramic to produce a pattern as shown in the enlarged region in Figure 13.
  • Other features of the core may also be provided to provide any required features of the resultant cooling passages, for example any required pattern to produce features such as the macro cooling feature 206 which interrupt the flow through the remainder of the cooling passage.
  • the micro grooves may be formed by machining the surface of the ceramic core.
  • the micro grooves may be formed by laser ablating the surface of the ceramic core.
  • the surface of the ceramic core may be laser ablated to remove small amounts of the ceramic, thereby forming a uniformly rough surface, i.e. a homogeneous region of surface treatment.
  • the surface roughness may be provided by applying a coarse coating to the ceramic core, where the coarse coating provides the desired predetermined surface roughness.
  • micro grooves are provided, they may be provided in a variety of patterns corresponding to the desired pattern of micro-ribs, as previously described. That is to say, the width, depth and orientation of the grooves should be provided such that a surface roughness (Ra) greater than 7 ⁇ m but no more 15 ⁇ m is achieved on the resultant component.
  • Ra surface roughness
  • a component for a turbo machine which may be provided as a rotor blade, with a cooling passage, a region of the cooling passage having a predetermined surface roughness in the range of about 7 ⁇ m but no more than about 12 ⁇ m to thereby enhance surface cooling in that region.
  • a method of manufacturing such a component is also provided.
  • the regions of predetermined surface roughness provide enhanced internal cooling to ensure that the metal temperatures are low enough to prevent excessive oxidation, and provide an adequate creep life.
  • the heat transfer can be significantly enhanced thus allowing greater engine efficiency, or a longer service life.
  • Provision of predetermined roughness according to the present disclosure is also advantageous as it can be use in addition with conventional turbulators (macro ribs) or impingement jets to further enhance the heat transfer.
  • turbulators macro ribs
  • impingement jets to further enhance the heat transfer.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP17189178.1A 2017-09-04 2017-09-04 Composant et procédé de fabrication associé Withdrawn EP3450683A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP17189178.1A EP3450683A1 (fr) 2017-09-04 2017-09-04 Composant et procédé de fabrication associé

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP17189178.1A EP3450683A1 (fr) 2017-09-04 2017-09-04 Composant et procédé de fabrication associé

Publications (1)

Publication Number Publication Date
EP3450683A1 true EP3450683A1 (fr) 2019-03-06

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3094032A1 (fr) * 2019-03-22 2020-09-25 Safran Aircraft Engines Aube de turbomachine d’aeronef et son procede de fabrication par moulage a cire perdue
RU2800747C2 (ru) * 2019-03-22 2023-07-27 Сафран Эркрафт Энджинз Лопатка авиационного газотурбинного двигателя и способ ее изготовления посредством литья по выплавляемым восковым моделям

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5975850A (en) * 1996-12-23 1999-11-02 General Electric Company Turbulated cooling passages for turbine blades
GB2441148A (en) * 2006-08-23 2008-02-27 Rolls Royce Plc Gas turbine engine component with coolant passages
EP2602439A1 (fr) * 2011-11-21 2013-06-12 Siemens Aktiengesellschaft Composant de gaz chaud pouvant être refroidi pour une turbine à gaz
US20130280092A1 (en) * 2012-04-24 2013-10-24 Jinquan Xu Airfoil cooling enhancement and method of making the same
EP2947274A1 (fr) * 2014-05-22 2015-11-25 United Technologies Corporation Structures de refroidissement des opérations de mise en turbulence
EP2975351A1 (fr) * 2014-07-14 2016-01-20 United Technologies Corporation Finition de surface par fabrication additive
EP3168535A1 (fr) * 2015-11-13 2017-05-17 General Electric Technology GmbH Corps de forme aérodynamique et procédé de refroidissement d'un corps placé dans un écoulement de fluide chaud

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5975850A (en) * 1996-12-23 1999-11-02 General Electric Company Turbulated cooling passages for turbine blades
GB2441148A (en) * 2006-08-23 2008-02-27 Rolls Royce Plc Gas turbine engine component with coolant passages
EP2602439A1 (fr) * 2011-11-21 2013-06-12 Siemens Aktiengesellschaft Composant de gaz chaud pouvant être refroidi pour une turbine à gaz
US20130280092A1 (en) * 2012-04-24 2013-10-24 Jinquan Xu Airfoil cooling enhancement and method of making the same
EP2947274A1 (fr) * 2014-05-22 2015-11-25 United Technologies Corporation Structures de refroidissement des opérations de mise en turbulence
EP2975351A1 (fr) * 2014-07-14 2016-01-20 United Technologies Corporation Finition de surface par fabrication additive
EP3168535A1 (fr) * 2015-11-13 2017-05-17 General Electric Technology GmbH Corps de forme aérodynamique et procédé de refroidissement d'un corps placé dans un écoulement de fluide chaud

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3094032A1 (fr) * 2019-03-22 2020-09-25 Safran Aircraft Engines Aube de turbomachine d’aeronef et son procede de fabrication par moulage a cire perdue
WO2020193899A1 (fr) 2019-03-22 2020-10-01 Safran Aicraft Engines Aube de turbomachine d'aeronef et son procede de fabrication par moulage a cire perdue
CN113924406A (zh) * 2019-03-22 2022-01-11 赛峰飞机发动机公司 飞行器涡轮机叶片及使用失蜡铸造制造该叶片的方法
RU2800747C2 (ru) * 2019-03-22 2023-07-27 Сафран Эркрафт Энджинз Лопатка авиационного газотурбинного двигателя и способ ее изготовления посредством литья по выплавляемым восковым моделям

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