EP3495612B1 - Method for manufacturing composite component - Google Patents
Method for manufacturing composite component Download PDFInfo
- Publication number
- EP3495612B1 EP3495612B1 EP17205758.0A EP17205758A EP3495612B1 EP 3495612 B1 EP3495612 B1 EP 3495612B1 EP 17205758 A EP17205758 A EP 17205758A EP 3495612 B1 EP3495612 B1 EP 3495612B1
- Authority
- EP
- European Patent Office
- Prior art keywords
- parts
- metal
- manufacturing
- fibres
- metal matrix
- 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.)
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Links
- 238000004519 manufacturing process Methods 0.000 title claims description 28
- 238000000034 method Methods 0.000 title claims description 26
- 239000002131 composite material Substances 0.000 title claims description 18
- 239000002184 metal Substances 0.000 claims description 38
- 238000001816 cooling Methods 0.000 claims description 13
- 239000011159 matrix material Substances 0.000 claims description 12
- 239000000843 powder Substances 0.000 claims description 11
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims description 10
- 229910052799 carbon Inorganic materials 0.000 claims description 10
- 238000002844 melting Methods 0.000 claims description 9
- 230000008018 melting Effects 0.000 claims description 9
- 238000005516 engineering process Methods 0.000 claims description 4
- 238000004924 electrostatic deposition Methods 0.000 claims description 3
- 239000000463 material Substances 0.000 description 13
- 239000011156 metal matrix composite Substances 0.000 description 11
- 239000007789 gas Substances 0.000 description 8
- 238000005266 casting Methods 0.000 description 5
- 229910052755 nonmetal Inorganic materials 0.000 description 5
- 238000009718 spray deposition Methods 0.000 description 4
- 238000000151 deposition Methods 0.000 description 3
- 230000008021 deposition Effects 0.000 description 3
- 239000000835 fiber Substances 0.000 description 3
- 238000003754 machining Methods 0.000 description 3
- 230000003014 reinforcing effect Effects 0.000 description 3
- 238000003466 welding Methods 0.000 description 3
- 239000000654 additive Substances 0.000 description 2
- 230000000996 additive effect Effects 0.000 description 2
- 239000000919 ceramic Substances 0.000 description 2
- 238000007689 inspection Methods 0.000 description 2
- 238000005304 joining Methods 0.000 description 2
- 238000010276 construction Methods 0.000 description 1
- 239000012809 cooling fluid Substances 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 238000001514 detection method Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- 230000000977 initiatory effect Effects 0.000 description 1
- 229910001092 metal group alloy Inorganic materials 0.000 description 1
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical class C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 1
- 150000002843 nonmetals Chemical class 0.000 description 1
- 230000003647 oxidation Effects 0.000 description 1
- 238000007254 oxidation reaction Methods 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
- 239000012783 reinforcing fiber Substances 0.000 description 1
- 238000007669 thermal treatment Methods 0.000 description 1
Images
Classifications
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- 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
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- 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
- F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/02—Selection of particular materials
- F04D29/023—Selection of particular materials especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
- F04D29/324—Blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
-
- 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
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
- F05D2300/6032—Metal matrix composites [MMC]
Definitions
- the present invention relates to a composite component and a method for manufacturing the same.
- the composite component is used in a gas turbine; for example the composite component is a compressor blade or vane or a turbine blade or vane for a gas turbine.
- Metal matrix composite materials have metal or ceramic reinforcing fibers embedded in a metal matrix; metal matrix composite materials have high strength and, compared to non-metal matrix composite materials, they can typically operate in a wider range of temperatures, do not absorb moisture, have better electrical and thermal conductivity, are less sensitive to stress concentration, for example due to notching effects or impacts of foreign objects. Metal matrix composite materials have a lower density and higher specific strength than non-metal matrix composite materials, thus they allow light weight construction.
- US 2013/0 259 701 A1 discloses a reinforcing edge of a turbomachine blade having a reinforcing structure of three-dimensionally woven ceramic fibers and a metal or metal alloy matrix.
- Non-metal matrix composite materials are also known, like carbon fiber composites such as carbon-fiber-reinforced carbon (CFRC); these non-metal matrix composite materials can withstand high temperatures but they cannot be exposed to an oxidizing atmosphere at high temperatures, are sensitive to impact damages, their life time or failure is difficult to predict and current inspection technologies do not allow a prediction of the remaining life time.
- CFRRC carbon-fiber-reinforced carbon
- the inventors of the present description have found a way to combine the advantages of the metal matrix composite materials with those of the non-metal matrix composite materials.
- EP 2 474 638 A2 discloses a method for manufacturing a composite component for a gas turbine that comprises at least a first part made out of a metal matrix reinforced with carbon fibres. The first part is joined to at least a second part made out of metal without carbon fibres.
- EP 3 170 587 A2 teaches providing metal powder, providing carbon fibres, simultaneously manufacturing the first parts and the second parts by local laser melting technology, such as SLM or direct laser melting.
- SLM local laser melting technology
- Other examples of known methods are disclosed in US 6 144 008 A , in US 9 32 446 B2 , in EP 2 703 605 A2 , in EP 3 020 918 A1 and in EP 1 920 869 A1 .
- the present invention concerns a method for manufacturing a composite gas turbine blade or vane in accordance with the accompanying claims.
- a composite component 1 for a gas turbine is a compressor blade or vane or a turbine blade or vane; other components are possible, such as heat shields, etc, however such components do not fall under the scope of the invention
- the component 1 comprises at least a first part 2 made out of a metal matrix 3 reinforced with carbon fibres 4; the at least a first part 2 is joined to at least a second part 6 made out of metal 7 without reinforcing carbon fibres.
- the metal of the metal matrix 3 of the first part 2 can be different from the metal 7 of the second part 6.
- the metal of the first and second parts 2, 6 can be selected according to the properties desired, such as weight, erosion resistance, corrosion or oxidation resistance, etc.
- the metal of the metal matrix 3 of the first part 2 can be the same as the metal 7 of the second part 6. This solution can be preferred if the manufacturing process so requires or in case the first and second parts 2, 6 require metal having the same properties.
- Components exposed to high temperature gases of a gas turbine e.g. compressor or turbine blades or vanes
- Components made out of traditional metal matrix composite materials cannot be provided with cooling systems (at the current state of the art), therefore the application of traditional metal matrix composite materials in gas turbine components is limited due to the maximum temperature the metal matrix can withstand.
- one or more second parts 6 can comprise at least a cooling element 8.
- the cooling element 8 is preferably a channel, which can carry a cooling fluid, such as air.
- the turbine or compressor blade or vane has a nose 9 that is defined by a second part 6 and opposite sides that are defined by first parts 2.
- An intermediate section of the component 1 can be defined by an additional second part 6 and the terminal part with the trailing edge can be defined by a first part 2, e.g. to be able to manufacture a thin but at the same time strong trailing edge.
- the composite component 1 can withstand high temperatures, because the first parts 2 include fibres able to withstand high stress also at high temperature.
- these first parts can be cooled by the second parts that are provided with the cooling elements 8.
- a first example of the method for manufacturing the composite component which does not fall under the scope of the claims comprises ( figures 2 through 3 ): manufacturing at least a first part 2 and a second part 6, then joining the at least a first part 2 and second part 6.
- Manufacturing of the first and second parts 2, 6 according to this embodiment of the method occurs separately, i.e. the first parts 2 are manufactured separately from the second parts 6; this allows to advantageously select the best method for manufacturing each part 2 or 6, according to the required features thereof.
- possible manufacturing methods for the first and/or second parts are casting, additive manufacturing such as SLM (selective laser melting), machining, spray deposition, etc.
- Joining the first part 2 to the second part 6 comprises laser welding or laser deposition welding.
- first parts 2 are manufactured by casting or SLM or spray deposition of metal on the carbon fibres.
- second parts 6 are casted or are manufactured by SLM or are manufactured by spray deposition.
- the first parts 2 and second parts 6 are then welded together (reference 10 identifies the welding).
- a second example of the method for manufacturing the composite component which does not fall under the scope of the claims comprises:
- this method allows manufacturing of first parts 2 whose features (e.g. material or geometrical features or manufacturing method or mechanical/thermal treatments) are independent from those of the second parts 6, because the first parts 2 are built before and thus independently of the second parts 6.
- features e.g. material or geometrical features or manufacturing method or mechanical/thermal treatments
- first parts 2 when the first parts 2 are manufactured, they can be provided with fibres that protrude from them ( figure 5 ), such that when the first parts 2 are housed in the mold 11 and the second parts 6 are casted, the protruding fibres promote holding of the first parts 2 to the second parts 6.
- first parts 2 are manufactured first, e.g. by casting, additive manufacturing such as SLM (selective laser melting), spray deposition, machining, etc.. Then these first parts 2 are housed in the mold 11 where metal is introduced to manufacture also the second parts 6 directly joined to the first parts 2.
- additive manufacturing such as SLM (selective laser melting), spray deposition, machining, etc.
- the cooling elements 8 can be made during casting of the second part 6; alternatively or in addition, the cooling elements 8 can be realized e.g. by machining or in other ways after casting of the second parts 6.
- a third example of the method for manufacturing the composite component 1 which does not fall under the scope of the claims comprises:
- Bonding of the first part 2 to the second parts 6 is in this embodiment particularly effective, because all parts 2, 6 are made at the same time in the mold 11.
- the fibre structure 16 makes it easier and faster handling of the fibres.
- the method for manufacturing the composite component comprises:
- This method allows manufacturing of complex three dimensional shapes with high tensile strength.
- the design of the first parts 2 can be easily optimized as load carrying sections and the design of the second parts 6 can be easily optimized in view of the required cooling.
- the powder in order to carry out the SLM process, can be deposited by electrostatic deposition, this allows to easily fill in gaps between fibres and to also build in vertical or inclined or downwards direction.
- the angle A of the laser beam 18 with the support plane 19 of the fibres is between 10-90 degree and preferably between 30-70 degree, with a thickness of the metal powder layer h less than 0.5 times the diameter D of the fibres and preferably the thickness of the metal powder layer is less than 0.3 times the diameter D of the fibres.
- the distance d between the fibres is greater than 0.6 times the diameter D of the fibres and preferably 0.8 times and is less than 2 times the diameter D of the fibres and preferably 1.2 times.
- the diameter DL of the laser beam is greater than 1.1 times the diameter D of the fibres and preferably it is 1.5 times greater than the diameter D, such that the laser beam heats the fibres and extends beyond their sides to melt the metal powder.
- the advantages of an impact resistant metal matrix are combined with light weight temperature resistant fibers.
- the surface of the metal matrix allows inspection and detection of crack initiation.
- the leading edge can be defined by a first part 2 (to allow for intense cooling via the cooling element 8) and the trailing edge can be defined by a second part 6, to be able to provide a strong, thin trailing edge.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Materials Engineering (AREA)
- Composite Materials (AREA)
- Architecture (AREA)
- Powder Metallurgy (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
- The present invention relates to a composite component and a method for manufacturing the same. In particular the composite component is used in a gas turbine; for example the composite component is a compressor blade or vane or a turbine blade or vane for a gas turbine.
- Metal matrix composite materials have metal or ceramic reinforcing fibers embedded in a metal matrix; metal matrix composite materials have high strength and, compared to non-metal matrix composite materials, they can typically operate in a wider range of temperatures, do not absorb moisture, have better electrical and thermal conductivity, are less sensitive to stress concentration, for example due to notching effects or impacts of foreign objects. Metal matrix composite materials have a lower density and higher specific strength than non-metal matrix composite materials, thus they allow light weight construction.
US 2013/0 259 701 A1 discloses a reinforcing edge of a turbomachine blade having a reinforcing structure of three-dimensionally woven ceramic fibers and a metal or metal alloy matrix. - Non-metal matrix composite materials are also known, like carbon fiber composites such as carbon-fiber-reinforced carbon (CFRC); these non-metal matrix composite materials can withstand high temperatures but they cannot be exposed to an oxidizing atmosphere at high temperatures, are sensitive to impact damages, their life time or failure is difficult to predict and current inspection technologies do not allow a prediction of the remaining life time.
- The inventors of the present description have found a way to combine the advantages of the metal matrix composite materials with those of the non-metal matrix composite materials.
-
EP 2 474 638 A2EP 3 170 587 A2 teaches providing metal powder, providing carbon fibres, simultaneously manufacturing the first parts and the second parts by local laser melting technology, such as SLM or direct laser melting. Other examples of known methods are disclosed inUS 6 144 008 A , inUS 9 32 446 B2 , inEP 2 703 605 A2EP 3 020 918 A1EP 1 920 869 A1 - The present invention concerns a method for manufacturing a composite gas turbine blade or vane in accordance with the accompanying claims.
- Further characteristics and advantages will be more apparent from the description of a preferred but non-exclusive embodiment of the composite component and method, illustrated by way of non-limiting example in the accompanying drawings, in which:
-
Figure 1 shows the composite component; -
Figures 2 through 3 show a first exemplary embodiment of the method: -
Figures 4 through 8 show a second exemplary - embodiment of the method
-
Figures 9 through 12 show an embodiment of the method; -
Figures 13 through 14 show the SLM manufacturing in the area of the fibres. - With reference to the figures, these show a
composite component 1 for a gas turbine. The composite component is a compressor blade or vane or a turbine blade or vane; other components are possible, such as heat shields, etc, however such components do not fall under the scope of the invention - The
component 1 comprises at least afirst part 2 made out of ametal matrix 3 reinforced withcarbon fibres 4; the at least afirst part 2 is joined to at least asecond part 6 made out ofmetal 7 without reinforcing carbon fibres. - The metal of the
metal matrix 3 of thefirst part 2 can be different from themetal 7 of thesecond part 6. In this case the metal of the first andsecond parts - Alternatively, the metal of the
metal matrix 3 of thefirst part 2 can be the same as themetal 7 of thesecond part 6. This solution can be preferred if the manufacturing process so requires or in case the first andsecond parts - Components exposed to high temperature gases of a gas turbine (e.g. compressor or turbine blades or vanes) often require large amounts of cooling air, to control their temperature. Components made out of traditional metal matrix composite materials cannot be provided with cooling systems (at the current state of the art), therefore the application of traditional metal matrix composite materials in gas turbine components is limited due to the maximum temperature the metal matrix can withstand.
- Advantageously, in order to improve cooling of the
component 1, one or more second parts 6 (in case thecomponent 1 is provided with more than one second part, like in the example shown infigure 1 ) can comprise at least acooling element 8. - The
cooling element 8 is preferably a channel, which can carry a cooling fluid, such as air. - As shown in
figure 1 , the turbine or compressor blade or vane has anose 9 that is defined by asecond part 6 and opposite sides that are defined byfirst parts 2. An intermediate section of thecomponent 1 can be defined by an additionalsecond part 6 and the terminal part with the trailing edge can be defined by afirst part 2, e.g. to be able to manufacture a thin but at the same time strong trailing edge. - Advantageously, during operation the
composite component 1 can withstand high temperatures, because thefirst parts 2 include fibres able to withstand high stress also at high temperature. In addition these first parts can be cooled by the second parts that are provided with thecooling elements 8. - In order to manufacture the composite component described above different embodiments of the method are possible.
- A first example of the method for manufacturing the composite component which does not fall under the scope of the claims comprises (
figures 2 through 3 ):
manufacturing at least afirst part 2 and asecond part 6, then joining the at least afirst part 2 andsecond part 6. - Manufacturing of the first and
second parts first parts 2 are manufactured separately from thesecond parts 6; this allows to advantageously select the best method for manufacturing eachpart - Joining the
first part 2 to thesecond part 6 comprises laser welding or laser deposition welding. - For example three
first parts 2 are manufactured by casting or SLM or spray deposition of metal on the carbon fibres. Separately twosecond parts 6 are casted or are manufactured by SLM or are manufactured by spray deposition. Thefirst parts 2 andsecond parts 6 are then welded together (reference 10 identifies the welding). - A second example of the method for manufacturing the composite component which does not fall under the scope of the claims comprises:
- manufacturing the
first parts 2, - providing the
first parts 2 in amold 11, - providing metal into the
mold 11 to cast thesecond parts 6 joined to thefirst parts 2. - Advantageously, this method allows manufacturing of
first parts 2 whose features (e.g. material or geometrical features or manufacturing method or mechanical/thermal treatments) are independent from those of thesecond parts 6, because thefirst parts 2 are built before and thus independently of thesecond parts 6. - In addition, when the
first parts 2 are manufactured, they can be provided with fibres that protrude from them (figure 5 ), such that when thefirst parts 2 are housed in themold 11 and thesecond parts 6 are casted, the protruding fibres promote holding of thefirst parts 2 to thesecond parts 6. - For example, three
first parts 2 are manufactured first, e.g. by casting, additive manufacturing such as SLM (selective laser melting), spray deposition, machining, etc.. Then thesefirst parts 2 are housed in themold 11 where metal is introduced to manufacture also thesecond parts 6 directly joined to thefirst parts 2. - The
cooling elements 8 can be made during casting of thesecond part 6; alternatively or in addition, thecooling elements 8 can be realized e.g. by machining or in other ways after casting of thesecond parts 6. - A third example of the method for manufacturing the
composite component 1 which does not fall under the scope of the claims comprises: - providing at least a
prefabricate structure 16 made offibres 4 joined atpoints 17, - providing the
structure 16 into amold 11, - providing a metal into the
mold 11 to simultaneously cast thefirst parts 2 and thesecond parts 6; thesefirst parts 2 andsecond parts 6 are in this way realized already joined. - Bonding of the
first part 2 to thesecond parts 6 is in this embodiment particularly effective, because allparts mold 11. In addition, thefibre structure 16 makes it easier and faster handling of the fibres. - According to the invention, the method for manufacturing the composite component comprises:
- providing metal powder,
- providing
fibres 4, - simultaneously manufacturing the
first parts 2 and thesecond parts 4 by local laser melting technology, such as SLM or direct laser melting, - This method allows manufacturing of complex three dimensional shapes with high tensile strength. In addition, the design of the
first parts 2 can be easily optimized as load carrying sections and the design of thesecond parts 6 can be easily optimized in view of the required cooling. - Advantageously, in order to carry out the SLM process, the powder can be deposited by electrostatic deposition, this allows to easily fill in gaps between fibres and to also build in vertical or inclined or downwards direction.
- Advantageously, with direct laser melting, powder deposition along different directions is possible in order to allow deposition between adjacent fibres.
- When in any of the methods above the component is realized by SLM (with metal powder deposited by electrostatic deposition or in other ways), in order to correctly melt the metal powder also below the fibres (with reference to the manufacturing of the first parts 2), the angle A of the
laser beam 18 with thesupport plane 19 of the fibres is between 10-90 degree and preferably between 30-70 degree, with a thickness of the metal powder layer h less than 0.5 times the diameter D of the fibres and preferably the thickness of the metal powder layer is less than 0.3 times the diameter D of the fibres. The distance d between the fibres is greater than 0.6 times the diameter D of the fibres and preferably 0.8 times and is less than 2 times the diameter D of the fibres and preferably 1.2 times. The diameter DL of the laser beam is greater than 1.1 times the diameter D of the fibres and preferably it is 1.5 times greater than the diameter D, such that the laser beam heats the fibres and extends beyond their sides to melt the metal powder. - According to the present description the advantages of an impact resistant metal matrix are combined with light weight temperature resistant fibers. Further, the surface of the metal matrix allows inspection and detection of crack initiation. E.g. the leading edge can be defined by a first part 2 (to allow for intense cooling via the cooling element 8) and the trailing edge can be defined by a
second part 6, to be able to provide a strong, thin trailing edge. - In practice the materials used and the dimensions can be chosen at will according to requirements and to the state of the art.
Claims (4)
- A method for manufacturing a composite gas turbine blade or vane (1), the composite gas turbine blade or vane (1) comprising first parts (2) made out of a metal matrix (3) reinforced with carbon fibres (4), the first parts being joined to second parts (6) made out of metal (7) without carbon fibres, the method comprising:providing metal powder,providing carbon fibres (4), simultaneously manufacturing the first parts (2) and the second parts (6) by local laser melting technology, such as SLM or direct laser melting,wherein the gas turbine blade or vane (1) has a nose (9) that is defined by one of said second parts (6) comprising a cooling channel (8), opposite sides defined by said first parts (2), an intermediate section defined by one of said second parts (6) comprising a cooling channel (8) and a terminal part comprising the trailing edge is defined by one of said first parts (2).
- The method of claim 1, characterized in that the metal of the metal matrix (3) of the at least a first part (2) is different from the metal (7) of the at least a second part (6).
- The method of claim 1, characterized in that the metal of the metal matrix (3) of the at least a first part (2) is the same as the metal (7) of the at least a second part (6).
- The method of any one of the preceding claims, characterized in that in order to carry out the SLM process, the powder is deposited by electrostatic deposition.
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP17205758.0A EP3495612B1 (en) | 2017-12-06 | 2017-12-06 | Method for manufacturing composite component |
CN201811487079.XA CN109877318B (en) | 2017-12-06 | 2018-12-06 | Composite component and method for producing same |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP17205758.0A EP3495612B1 (en) | 2017-12-06 | 2017-12-06 | Method for manufacturing composite component |
Publications (2)
Publication Number | Publication Date |
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EP3495612A1 EP3495612A1 (en) | 2019-06-12 |
EP3495612B1 true EP3495612B1 (en) | 2021-05-12 |
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Family Applications (1)
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EP17205758.0A Active EP3495612B1 (en) | 2017-12-06 | 2017-12-06 | Method for manufacturing composite component |
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EP (1) | EP3495612B1 (en) |
CN (1) | CN109877318B (en) |
Families Citing this family (2)
Publication number | Priority date | Publication date | Assignee | Title |
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GB2582148A (en) * | 2019-03-12 | 2020-09-16 | Airbus Operations Ltd | Impact resistant panels |
US11732586B2 (en) * | 2020-05-14 | 2023-08-22 | Toyota Motor Engineering & Manufacturing North America, Inc. | Metal matrix composite turbine rotor components |
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US10370971B2 (en) * | 2014-11-17 | 2019-08-06 | United Technologies Corporation | Reinforced gas turbine engine rotor disk |
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2017
- 2017-12-06 EP EP17205758.0A patent/EP3495612B1/en active Active
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- 2018-12-06 CN CN201811487079.XA patent/CN109877318B/en active Active
Patent Citations (3)
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US6144008A (en) * | 1996-11-22 | 2000-11-07 | Rabinovich; Joshua E. | Rapid manufacturing system for metal, metal matrix composite materials and ceramics |
US9327446B2 (en) * | 2012-12-10 | 2016-05-03 | Rolls-Royce Plc | Joint structure and method |
EP3170587A2 (en) * | 2015-10-28 | 2017-05-24 | Airbus Operations GmbH | Fibre-reinforced metal component for an aircraft or spacecraft and production methods for fibre-reinforced metal components |
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CN109877318B (en) | 2023-08-04 |
CN109877318A (en) | 2019-06-14 |
EP3495612A1 (en) | 2019-06-12 |
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