• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a method for manufacturing a two-component blade for a gas turbine engine, comprising successively obtaining a blade profile (24) of ceramic material comprising a hole through the blade profile of partly in the longitudinal direction to form a longitudinal channel (28) opening into an upper cavity (26a), positioning and maintaining the blade profile in a mold (30) so as to provide a lower cavity ( 26b) communicating with the channel (28) of the blade profile, casting molten metal in the blade profile so as to fill the upper and lower cavities and the channel connecting them, and cooling the metal so the removal of the cooled metal in the upper and lower cavities causes compression prestressing of the ceramic of the blade profile.
    公开号:FR3023317A1
    申请号:FR1456426
    申请日:2014-07-04
    公开日:2016-01-08
    发明作者:Francois Casteilla；Brou De Cuissart Sebastien Digard；Serge Fargeas；Vincent Herb；Chantal Sylvette Langlois
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION The present invention relates to the general field of manufacture of gas turbine engine blades.
    [0002] The vanes of a gas turbine engine, and in particular the vanes of the high-pressure turbine of a turbojet, are subjected to the hot gases from the combustion chamber. To improve the efficiency of the engine, it is known to increase the temperature of these ejection gases from the combustion chamber to temperatures that may be much higher than the melting points of the best metal alloys in which the blades are generally manufactured . Also, to allow the metal alloy blades to hold at very high temperatures, it is known to cover them with a ceramic coating forming a thermal barrier and to practice internal cooling circuits. However, such coatings and such internal cooling circuits are increasingly complex to achieve and do not always allow the blades to hold at the very high temperature ejection gases from the combustion chamber. In particular, the temperature of these ejection gases is caused to increase beyond the current temperatures of the ejection gas from the combustion chamber, which further complicates the production of thermal barrier coatings and internal cooling circuits.
    [0003] OBJECT AND SUMMARY OF THE INVENTION There is therefore a need to be able to have a method of manufacturing a blade enabling it to withstand very high temperatures without presenting the aforementioned drawbacks.
    [0004] According to the invention, this object is achieved by a method for manufacturing a two-component blade for a gas turbine engine, comprising successively: obtaining a ceramic material blade profile comprising a through-hole the blade profile from side to side in the direction of its length to form a longitudinal channel opening at a first longitudinal end of the blade profile in an upper cavity; positioning and maintaining the blade profile in a mold so as to provide a lower cavity communicating with the blade profile channel at a second longitudinal end of the blade profile; casting molten metal in the blade profile so as to fill the upper and lower cavities and the channel connecting them; and cooling the metal so that removal of the cooled metal in the upper and lower cavities causes compressive prestressing of the ceramic of the blade profile. The manufacturing method according to the invention is remarkable in that, during the cooling of the cast metal in the cavities of the blade profile, it, by its natural shrinkage, will exert a compressive force on the ceramic constituting this profile dawn (in the direction of the length of the blade profile). The blade obtained by this manufacturing method therefore has a prestress in compression of the ceramic. In operation, the force experienced by the blade is a centrifugal force (in the direction of the length of the blade profile) which results in a tensile force on the components of the blade. With its prestress in compression and thanks to the expansion gap between the ceramic and the metal that constitute it, the blade will easily withstand the tensile forces it undergoes. In particular, to cancel this prestressing in compression, it would be necessary to reach temperatures for which the metal used no longer has mechanical strength, such temperatures being practically never reached in a gas turbine engine. In addition, compressing the ceramic blade profile constituting the blade obtained by the manufacturing method according to the invention allows to have a wide choice of ceramics, including less expensive ceramics than those usually used. The weight of the blade thus obtained is also less than for blades known from the prior art. Finally, such a blade is easily repairable, by performing a simple replacement of the ceramic. Preferably, the ceramic material used to form the blade profile is alumina and the metal used for casting is a nickel-based metal alloy.
    [0005] Also preferably, the method further comprises producing an internal cooling circuit of the blade. In this case, the realization of an internal cooling circuit of the blade may comprise, prior to the casting step of molten metal, the arrangement in the channel of at least one elongate core traversing from one side to the other the blade profile in the direction of its length, and the extraction of said core after the casting step of molten metal to form a passage of air passing through the blade from side to side. The core for producing the internal cooling circuit of the blade can be made of silica.
    [0006] The realization of such a cooling circuit is easy, particularly with respect to the internal cooling circuits of the blades known from the prior art. In addition, this cooling circuit has a low impact on engine performance. The invention also relates to a blade for a gas turbine engine which is obtained by the method as defined above. The invention further relates to a gas turbine engine comprising at least one such blade. BRIEF DESCRIPTION OF THE DRAWINGS Other features and advantages of the present invention will be apparent from the description given below, with reference to the accompanying drawings which illustrate embodiments having no limiting character. In the figures: FIG. 1 is a diagrammatic perspective view of a blade obtained by the manufacturing method according to the invention; FIGS. 2 to 5 are diagrammatic views showing different steps of the manufacturing process of the blade of FIG. 1; and - Figures 6 and 7 are longitudinal sectional views of blades according to alternative embodiments of the invention.
    [0007] DETAILED DESCRIPTION OF THE INVENTION The invention applies to the production of any blade fitted to a gas turbine engine, and in particular to the blades of the high-pressure turbine of a turbojet engine, such as the blade 10 shown in FIG. figure 35 1.
    [0008] 302 3 3 1 7 4 In a manner known per se, the blade 10 has a longitudinal axis XX and is intended to be fixed on a rotor disc of the high-pressure turbine of the turbojet engine via a fitting 12 generally fir-shaped. The blade 10 extends along the longitudinal axis XX between a foot 14 and a top 16 and comprises a leading edge 18 and a trailing edge 20. The fitting 12 connects to the foot 14 of the blade at level of a platform 22 defining an inner wall for the flow of the flue gas stream passing through the high pressure turbine.
    [0009] Such a blade 10 must withstand the very high temperatures of the gases coming from the combustion chamber of the turbojet engine located directly upstream of the high pressure turbine. According to the invention, there is provided a method of manufacturing such a blade that achieves this goal.
    [0010] For this purpose, the method according to the invention initially provides for the production of a blade profile of ceramic material. By dawn profile, we mean a piece having the final shape of the dawn. This blade profile can be obtained according to various techniques known per se and not described in detail here, for example by a method 20 implementing the injection of ceramic into a mold of appropriate shape or by a manufacturing process called "additive "(Ie using three-dimensional printing). Another known technique that can be used (suitable for mass production of dawn) is the lost wax casting process with a shell mold and using directed solidification. Reference can be made to EP 2,092,999 in which such a method is described. When the process used to obtain the blade profile involves the injection of ceramic into a mold of appropriate shape, the blade profile is then drilled (for example by a mechanical tool) to make a through hole the blade profile from side to side in the direction of its length, that is to say along the longitudinal axis XX of the blade to manufacture. This through hole is made to form an upper cavity 26a at an upper longitudinal end of the blade profile, this upper cavity opening into a channel 28. The channel 28 has a smaller diameter d28. dimension that the upper cavity 26a in which it opens. As shown in FIG. 3, the blade profile 24 is positioned and held in place in a mold 30 so as to provide a lower cavity 26b communicating with the channel 28 of the blade profile at the lower longitudinal end. of the dawn profile. As for the upper cavity, this lower cavity 26b has a cross-section of larger dimension than the diameter d28 of the channel 28. Of course, when the method used to obtain the blade profile uses other techniques manufacturing, such as in particular lost wax foundry with a shell mold, the channel and the upper and lower cavities of the blade profile will be obtained by other means (typically by cores). The next step of the process according to the invention consists in producing a casting of molten metal in the blade profile 24 so as to fill the two cavities 26a, 26b and the channel 28 which connects them. This casting operation is typically carried out in a casting furnace (not shown in the figures) inside which the mold 30 is placed. The molten metal Mf is poured into the mold 30 through the blade profile 24 from the upper cavity 26a to completely fill the lower cavity 26b, the channel 28 and the upper cavity 26a. By way of example, a metal alloy based on nickel, such as AMI, may be chosen as the casting metal. typically used by Snecma for the production of some of its monocrystalline turbine blades. When the casting of the molten metal is finished, the mold 30 is taken out of the casting furnace and the blade profile 24 is cooled. This cooling has the natural consequence of causing a withdrawal of the metal inside the cavities 26a, 26b and the channel 28 which connects these last 30 (this withdrawal is symbolized by the arrows in Figure 4). Thus, as shown in FIG. 5, a two-component vane 10 is obtained with the metal-filled upper cavity 26a situated on the side of the apex 16 of the vane and the lower cavity 26b defining at least in part the shank 12 of the dawn.
    [0011] The removal of the cooled metal in the cavities and the channel which connects them causes compression prestressing of the ceramic 302 3 3 1 7 6 of the blade profile (this compressive prestressing is symbolized by the arrows Fc). In operation, the blade is subjected, on the one hand to high temperatures, and on the other hand to a centrifugal force (in the direction of the length of the blade profile, of the fitting towards the top), this centrifugal force resulting in a tensile force on the blade profile. With its prestress in compression, the blade can easily withstand the tensile forces it undergoes. In particular, the expansion gap between the ceramic constituting the blade profile and the metal cast in the cavities and the channel thereof decreases in operation the compressive prestressing of the ceramic blade profile. FIG. 6 represents an embodiment variant of a blade 10 'obtained by the process according to the invention. In this variant, it is intended to provide an internal cooling circuit of the blade. For this purpose, prior to the casting step of molten metal in the blade profile, an elongated core (not shown), for example made of removable silica, is disposed in the channel 28 connecting the two cavities 26a, 26b between them. This core thus crosses right through the blade profile in the direction of its length.
    [0012] After the casting step of the molten metal, this core is extracted by unsticking the blade profile so as to form an air passage 30 passing through the blade 10 'from one side to the other. A flow of cooling air is introduced into this air passage 30 from the base of the blade fitting 12 and is discharged into the vein of the high pressure turbine at the top 16 of the dawn. . 7 represents another variant embodiment of a blade 10 "obtained by the method according to the invention In this other variant, the blade 10" also has an internal cooling circuit. When the process used to obtain the blade profile involves the injection of ceramic into a mold of appropriate shape, this circuit is formed, prior to the step of casting molten metal in the profile of dawn or when setting the blade profile in the casting mold, positioning a bush-shaped core in the channel 28. When the method used to obtain the blade profile implements the lost-wax casting with a shell mold, the cooling circuit is formed before the wax injection step by positioning, in the wax injection mold, a core of shape corresponding to the internal cooling circuit . After the casting step of the molten metal, this sleeve-shaped core is extracted from the blade profile so as to form an air passage 30 'which is annular in the central part of the blade 10'. cooling air flow is introduced into this air passage 30 'from the base of the fitting 12 of the blade and is discharged into the vein of the high-pressure turbine at the top 16 of the blade.
    权利要求:
    Claims (7)
    [0001]
    REVENDICATIONS1. A method of manufacturing a blade (10; 10 '; 10 ") two-component for a gas turbine engine, comprising successively: obtaining a blade profile (24) of ceramic material comprising a hole through the profile of vane through its longitudinal direction to form a longitudinal channel (28) opening at a first longitudinal end of the vane profile in an upper cavity (26a); positioning and maintaining the vane profile in a mold (30) for providing a lower cavity (26b) communicating with the channel (28) of the blade profile at a second longitudinal end of the blade profile; molten metal casting (Mf) in the vane profile so as to fill the upper and lower cavities and the channel connecting them, and the cooling of the metal so that the removal of the cooled metal in the upper and lower cavities causes compression prestressing ceramic of the dawn profile.
    [0002]
    The method of claim 1, wherein the ceramic material used to form the blade profile is alumina and the metal used for casting is a nickel-based metal alloy.
    [0003]
    3. Method according to claim 1 or 2, further comprising producing an internal cooling circuit of the blade.
    [0004]
    4. The method of claim 3, wherein the realization of an internal cooling circuit of the blade comprises, prior to the casting step of molten metal, the arrangement in the channel (28) of at least one an elongate core passing right through the blade profile along its length, and extracting said core after the molten metal casting step to form an air passage (30; 30 ') passing through the dawn from one side to the other.
    [0005]
    5. Method according to claim 4, wherein the core for producing the internal cooling circuit of the blade is made of silica.
    [0006]
    6. A blade (10; 10 '; 10 ") for a gas turbine engine, characterized in that it is obtained by the method according to any one of claims 1 to 5.
    [0007]
    7. Gas turbine engine comprising at least one blade according to claim 6.
    类似技术:
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    同族专利:
    公开号 | 公开日
    CA2954024A1|2016-01-07|
    WO2016001544A1|2016-01-07|
    CN106536089B|2019-05-03|
    BR112017000100B1|2021-09-14|
    RU2017103636A|2018-08-06|
    RU2687949C2|2019-05-16|
    RU2017103636A3|2018-11-28|
    FR3023317B1|2016-08-05|
    JP6741647B2|2020-08-19|
    US20170136534A1|2017-05-18|
    EP3164237B1|2019-10-02|
    BR112017000100A2|2017-10-31|
    EP3164237A1|2017-05-10|
    JP2017532475A|2017-11-02|
    US10486230B2|2019-11-26|
    CN106536089A|2017-03-22|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3844728A|1968-03-20|1974-10-29|United Aircraft Corp|Gas contacting element leading edge and trailing edge insert|
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    FR2463849A1|1979-08-23|1981-02-27|Onera |Blade for gas turbine rotor - has outer ceramic liner fitted over metal core and held by enlarged head and pin into rotor root fixing|
    GB2062530A|1979-11-10|1981-05-28|Mtu Muenchen Gmbh|Blade for rotary machines such as turbines|
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    EP2468434A1|2010-12-23|2012-06-27|General Electric Company|Processes for producing components containing ceramic-based and metallic materials|
    GB191300496A|1912-06-22|1913-04-03|Emil Spiegel|Improved Manufacture of Wheel Tyres.|
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    US6913064B2|2003-10-15|2005-07-05|United Technologies Corporation|Refractory metal core|
    FR2927270B1|2008-02-08|2010-10-22|Snecma|PROCESS FOR MANUFACTURING DIRECTED SOLIDIFICATION AUBES|US10697305B2|2016-01-08|2020-06-30|General Electric Company|Method for making hybrid ceramic/metal, ceramic/ceramic body by using 3D printing process|
    US11135031B2|2018-06-15|2021-10-05|Verb Surgical Inc.|User interface device having grip linkages|
    US20200102852A1|2018-09-28|2020-04-02|General Electric Company|Metallic compliant tip fan blade|
    US10920607B2|2018-09-28|2021-02-16|General Electric Company|Metallic compliant tip fan blade|
    法律状态:
    2015-08-10| PLFP| Fee payment|Year of fee payment: 2 |
    2016-01-08| PLSC| Publication of the preliminary search report|Effective date: 20160108 |
    2016-08-04| PLFP| Fee payment|Year of fee payment: 3 |
    2017-05-02| PLFP| Fee payment|Year of fee payment: 4 |
    2018-06-21| PLFP| Fee payment|Year of fee payment: 5 |
    2018-06-29| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170719 |
    2020-06-23| PLFP| Fee payment|Year of fee payment: 7 |
    2021-06-23| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1456426A|FR3023317B1|2014-07-04|2014-07-04|METHOD FOR MANUFACTURING A BI-COMPONENT BLADE FOR A GAS-TURBINE ENGINE OBTAINED BY SUCH A METHOD|FR1456426A| FR3023317B1|2014-07-04|2014-07-04|METHOD FOR MANUFACTURING A BI-COMPONENT BLADE FOR A GAS-TURBINE ENGINE OBTAINED BY SUCH A METHOD|
    BR112017000100-4A| BR112017000100B1|2014-07-04|2015-06-29|PROCESS FOR MANUFACTURING A TWO-COMPONENT BLADE FOR A GAS TURBINE ENGINE|
    RU2017103636A| RU2687949C2|2014-07-04|2015-06-29|Method of making two-component blade for gas turbine engine and blade, obtainable by such method|
    JP2017500065A| JP6741647B2|2014-07-04|2015-06-29|Method for manufacturing a two-component blade for a gas turbine engine and a blade obtained by the method|
    PCT/FR2015/051747| WO2016001544A1|2014-07-04|2015-06-29|Method for manufacturing a two-component blade for a gas turbine engine and blade obtained by such a method|
    CA2954024A| CA2954024A1|2014-07-04|2015-06-29|Method for manufacturing a two-component blade for a gas turbine engine and blade obtained by such a method|
    EP15736568.5A| EP3164237B1|2014-07-04|2015-06-29|Method for manufacturing a two-component blade for a gas turbine engine and blade obtained by such a method|
    CN201580036579.XA| CN106536089B|2014-07-04|2015-06-29|The blade that manufacture is used for the method for double component blades of gas-turbine unit and is obtained by this method|
    US15/323,766| US10486230B2|2014-07-04|2015-06-29|Method for manufacturing a two-component blade for a gas turbine engine and blade obtained by such a method|
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