WO2014053751A1

WO2014053751A1 - Method for producing an aerodynamic part by overmoulding a ceramic shell onto a composite preform

Info

Publication number: WO2014053751A1
Application number: PCT/FR2013/052302
Authority: WO
Inventors: Pascal Diss; Eric Bouillon; André Lafond; Michel Bourgeon
Original assignee: Herakles
Priority date: 2012-10-04
Filing date: 2013-09-27
Publication date: 2014-04-10
Also published as: FR2996549B1; FR2996549A1

Abstract

A method for producing an aerodynamic part comprising: - producing a preform (300) from a composite material made of a fibrous structure densified by a matrix, said preform having dimensions smaller than those of the aerodynamic part to be produced, - placing the preform made from composite material into a mould (400) having the same internal dimensions and shape as those of the aerodynamic part to be produced, - interposing a ceramic powder (50) between the preform made from composite material and the internal surface (402a) of the mould, said ceramic powder having a thermal expansion coefficient varying, at most, by 4.10"6.K_1 relative to the thermal expansion coefficient of the composite material of the preform, said ceramic powder further having a particle size of between 0.25 µm and 25 µm, - stabilising said powder in such a way as to form, around the preform made from composite material, a ceramic shell having the dimensions and the shape of the aerodynamic part to be produced.

Description

Method of manufacturing an aerodynamic part by overmoulding a ceramic envelope on a composite preform Background of the invention

The present invention relates to aerodynamic parts made of composite material comprising a fiber reinforcement densified by a matrix.

In aeronautical engines and in particular in the gas turbines of such engines, parts with aerodynamic shapes, such as blades, are usually made of metal alloys by a foundry process and local machining.

In order to significantly reduce the mass and to admit operating temperatures higher than those permitted with current metal alloys, a solution for producing the blades consists of using composite materials, such as, for example, ceramic matrix composite materials.

Indeed, the ceramic matrix composite materials (CMC) are part of so-called thermostructural composite materials, that is to say composite materials having good mechanical properties and ability to retain these properties at high temperature. In addition, parts, such as blades, made of CMC have a significant weight gain compared to the same parts made with the usual metal alloys.

In a well known manner, the CMC parts are formed by a fibrous reinforcement of refractory fibers (carbon or ceramic) which is densified by a ceramic matrix, in particular carbide, nitride, refractory oxide, .... Typical examples of CMC materials are C-SiC materials (carbon fiber reinforcement and silicon carbide matrix), SiC-SiC materials and CC / SiC materials (mixed carbon / silicon carbide matrix). The manufacture of CMC composite parts is well known.

However, CMC parts have a wavy and relatively rough surface appearance that may be incompatible with the aerodynamic performance required for parts such as blades. The surface corrugation is due to the fibrous reinforcement while the roughness is related to the ceramic matrix in "seal-coat", in particular when it is deposited by chemical vapor infiltration (CVI).

Conversely, parts made of metal alloys and associated processes have a smooth surface appearance with a very low roughness (of the order of 1 pm).

In order to improve the surface condition of the parts made of composite material, a first solution consists in carrying out a complete machining of the profile of the part in order to limit the ripple related to the texture. However, this operation only partially solves the problem of ripple and is not compatible with a series production of parts. In addition, the machining leads to the cutting of the reinforcing wire on the surface and consequently reduces the mechanical performance of the part.

Another solution for improving the surface condition of a composite material part consists in depositing on the surface thereof a coating fulfilling the following conditions:

a) adhesion to the composite material of the part,

b) compatibility with the coefficient of thermal expansion of the composite material of the part,

(c) stability under conditions of use,

d) formatting temperature compatible with the thermostability of the composite material of the part.

This solution is implemented by applying to the brush on the surface of the composite material part a suspension based on ceramic powder, for example SiC, and an organosilicon binder in order to form a ceramic coating on the surface of the piece. . However, if this implementation is compatible with the conditions set out above, its major disadvantage lies in the complexity of the formatting process. Indeed, the application being carried out with a brush, a sanding step is necessary after the crosslinking of the binder in order to erase the imperfections left during the application of the coating. This sanding step is particularly difficult on non-planar parts, requiring, at this stage, to perform the operation manually. In addition, in order to limit the formation of bubbles or cracks, only a fine deposit can be applied. Thus, to reach the desired surface state, it may be necessary to repeat the paint / crosslinking / sanding sequences (up to 5 times) several times. Finally, in order to ensure cohesion of the coating during its use, it is necessary to stabilize it, by performing, last, a step of densification by chemical vapor infiltration (CVI) of SiC which allows to link between them the ceramic grains. Such a CMC part surface treatment method is described in US 2006/0141154.

Although this method makes it possible to significantly improve the surface state of a composite material part by reducing the corrugations and the surface roughness, the complexity of its implementation causes a significant increase in the cost and the manufacturing time. of the room.

Obiet and summary of the invention

The present invention therefore aims to provide a method for manufacturing, industrially and economically, parts incorporating the mechanical properties of composite materials while respecting the requirements of shape and aerodynamic performance.

For this purpose, the invention proposes a method for manufacturing an aerodynamic part comprising:

the production of a preform made of composite material from a fibrous structure densified by a matrix, said preform having dimensions smaller than those of the aerodynamic part to be produced,

placing the preform made of composite material in a mold having internal dimensions and shape corresponding to those of the aerodynamic part to be produced,

the interposition of a ceramic powder between the preform made of composite material and the internal surface of the mold,

stabilizing said powder so as to form around the preform made of composite material a ceramic envelope with the dimensions and the shape of the aerodynamic part to be produced, a method in which the ceramic powder has a thermal expansion coefficient of not more than ± 4.10 ^"6 .K ^{" 1} with respect to the thermal expansion coefficient of the composite material of the preform, said ceramic powder further having a particle size of between 0 , 25 μιη and 25 Mm.

The method of the invention makes it possible to obtain an aerodynamic part which comprises both the necessary structural functions thanks to its preform made of composite material and the aerodynamic functions thanks to the controlled surface state of the ceramic envelope formed on the preform .

In addition, the overmoulding of the ceramic skin on the preform made of composite material of dimensions smaller than the final dimensions of the part to be produced makes it possible, in a minimum of steps that can be implemented industrially, to produce a so-called "near net shape" Of the piece, that is to say, obtaining, immediately after overmolding the ceramic casing, a piece to the shape and final dimensions required (no resumption of the necessary part).

Moreover, the use of a ceramic powder having a coefficient of thermal expansion which is sufficiently close to that of the composite material of the part as well as a particle size determined as above makes it possible to directly form the ceramic envelope on the material composite of the preform. Indeed, in addition to a suitable coefficient of thermal expansion, the ceramic powder has a particle size between a minimum value (0.25 pm) and a maximum value (25 pm). In the case of a stabilization of the ceramic powder by flash sintering, below the minimum value of determined particle size, premature re-agglomeration phenomena can occur and lead to the formation of particles of particle size that are too high to obtain ceramic shell with a satisfactory structure and surface finish. In the case of stabilization by chemical vapor infiltration of silicon carbide, if the powder has a particle size smaller than that determined in the invention, it becomes difficult for the infiltration gases to penetrate the powder, which risks to prevent a good stabilization of the latter. In addition, beyond the value of maximum particle size defined in the invention, the contact surface between the grains becomes too small to ensure a satisfactory cohesion of the ceramic envelope formed. In the case of a stabilization of the ceramic powder by flash sintering, a particle size that is too high requires the use of a thermal budget (temperature / time pair) for the sintering which is incompatible with the thermal stability temperature of the fibers of the composite material of FIG. the preform.

Thanks to the process of the invention, it is therefore possible to directly form a ceramic envelope on the composite material of the preform, which represents an important advantage, particularly in the case of the formation of such an envelope by overmoulding . Indeed, in order to ensure a good adhesion of the coating on the substrate, one or more attachment or adaptation sub-layers are generally used. However, in the case of overmolding, the use of one or more sub-layers is problematic because of the variability of the surface condition of the composite material preform. Indeed, there are risks of exposing the or the underlayments at the end of the overmolding operation because, in certain areas, the final thickness of the coating, which is between the outer surface of the preform and the inner surface of the mold, may be less than the thickness of the, or sub-layers, not allowing the formation of a ceramic envelope in these areas.

According to a first aspect of the process of the invention, the stabilization of the ceramic powder is carried out by flash sintering. In this case, a ceramic powder composition having a maximum flash sintering temperature greater than 250 ° C. is preferably chosen at the temperature of thermal stability of the fibers of the preform made of composite material.

According to a second aspect of the process of the invention, the interposition of the ceramic powder is obtained by mixing said ceramic powder with a binder and injecting the mixture between the preform and the inner surface of the mold, stabilizing the powder being made by chemical vapor infiltration of silicon carbide. The binder can be chosen in particular from at least one organosilicon compound and a fugitive resin.

In order to reinforce the mechanical strength of the ceramic envelope, it is also possible to have a mat of ceramic fibers around the fiber preform before the injection of the binder and ceramic powder mixture. Ceramic fibers may also be added to the binder and ceramic powder mixture.

According to a particular characteristic of the invention, the ceramic powder comprises nanopowders.

According to another particular characteristic of the invention, the preform is made of ceramic matrix composite material (CMC).

According to a third aspect of the invention, the aerodynamic part is a fixed or moving turbomachine blade.

Brief description of the drawings

Other characteristics and advantages of the invention will emerge from the following description of particular embodiments of the invention, given by way of non-limiting examples, with reference to the appended drawings, in which:

FIG. 1 is a perspective view of a turbomachine blade made from a process according to the invention,

FIG. 2 is a flowchart illustrating successive steps of modes of implementation of a method according to the invention,

FIG. 3 is a perspective view of the composite material preform of the blade of FIG. 1;

FIG. 4 is a three-dimensional view showing the surface state of a portion of a CMC part without additional surface treatment,

FIG. 5 is a curve for measuring the dimensional variations of the part portion of FIG. 4;

FIG. 6 is a curve for measuring the dimensional variations on the surface of a metallic material used for producing aeronautical motor blades, FIGS. 7A and 7B are perspective views showing the overmoulding of a ceramic envelope on the preform of FIG. 3 by flash sintering,

FIGS. 8A and 8B are photographs of a fiber preform on which a ceramic envelope has been overmolded according to a process of the invention,

FIG. 9 is a flowchart illustrating successive steps of modes of implementation of another method according to the invention,

FIGS. 10 and 11 show results obtained after the formation of a ceramic envelope made by overmolding a ceramic powder, the powder respectively having a particle size outside and within the range of particle size defined in FIG. invention,

FIGS. 12 and 13 show results obtained after the formation of a ceramic envelope made by overmolding a ceramic powder, the powder respectively having a particle size outside and within the range of particle size defined in FIG. invention

Detailed description of embodiments

The present invention provides a method for the manufacture of aerodynamic parts. The method of the invention relates in particular, but not exclusively, to the manufacture of blades for rotors or fixed wheels of gas turbines used in aircraft engines or industrial turbines. According to the invention and as described in detail below, each aerodynamic part is made from a composite material preform (fiber reinforcement densified by a matrix) having dimensions smaller than those of the part to be produced and covered by a ceramic envelope having the final dimensions of the aerodynamic part. The ceramic casing is obtained by molding and stabilizing / densifying a ceramic powder around the preform of the part. By "ceramic powder" is meant here a powder containing at least one or more constituents of the ceramic of the envelope to be formed on the preform. The ceramic powder may be further mixed with other elements in particular in the form of grains, liquid or fibers.

FIG. 1 illustrates a vane 100 of a low pressure turbine (BP) impeller which comprises a blade 120 and a foot 130 formed by a portion of greater thickness, for example with a bulbous section, extended by a stilt 132. The blade 120 extends longitudinally between its foot 130 and its top 121 and has a cross section a curved profile of variable thickness defining two faces 122 and 123, respectively corresponding to the extrados and intrados of the blade 120 and each connecting the leading edge 120a and the trailing edge 120b of the latter. In the example described here, the blade 120 further comprises a blade platform 140 and a blade root 160. In accordance with the invention, the blade 100 is formed of a preform 300 made of composite material, that is to say comprising a fiber reinforcement densified by a matrix, the preform 300 being further covered by a ceramic casing 150 defining the exact shape and dimensions of the blade.

Referring to Figure 2, a method of manufacturing the blade of Figure 1 according to the invention comprises the following steps.

The manufacture of the blade begins with the provision of a fibrous structure from which will be formed a composite material preform whose shape is close to that of the blade to be manufactured but whose dimensions are smaller than the final dimensions of the blade. dawn to achieve (step 10).

The fibrous structure can be in various forms, such as:

- two-dimensional fabric (2D),

- three-dimensional fabric (3D) obtained by 3D or multilayer weaving,

- braid,

- knit,

- felt,

- Unidirectional web (UD) of son or cables or multidirectional webs (nD) obtained by superposition of several UD webs in different directions and UD web connection between them for example by sewing, by chemical bonding agent or by needling. It is also possible to use a fibrous structure formed of several superimposed layers of fabric, braid, knit, felt, plies or others, which layers are bonded together, for example by sewing, by implantation of threads or rigid elements or by needling.

The fibers constituting the fibrous structure are refractory fibers, that is to say ceramic fibers, for example silicon carbide (SiC), carbon fibers or even fibers made of a refractory oxide, for example alumina (Al2O3).

In the example described here, the fibrous texture is produced by three-dimensional weaving or multi-layer in one piece of silicon carbide son such as in particular described in WO 2010/061140 and whose contents are incorporated herein by reference.

Once formed, the fibrous texture is shaped and consolidated by impregnation of the latter with a liquid composition containing a ceramic precursor consolidation resin (step 20).

For this purpose, the fibrous texture is immersed in a bath containing the resin and usually a solvent thereof. After draining, drying is carried out in an oven. The drying may be accompanied by a pre-crosslinking or partial crosslinking of the resin. Such pre-crosslinking providing additional stiffness, it must, if it is performed, remain limited to maintain sufficient deformability of the fibrous texture.

Other known impregnation techniques may be used such as preparing a prepreg by passing the fibrous texture in a continuous impregnator, impregnation by infusion, or impregnation by RTM ("Resin Transfer Molding").

The consolidation resin is chosen to leave, after pyrolysis, a ceramic residue sufficient to ensure consolidation of the fibrous preform made thereafter.

A ceramic precursor resin may be, for example, a polycarbosilane precursor resin of silicon carbide (SiC), or a polysiloxane resin precursor of SiCO, or a polyborocarbosilazane resin precursor of SiCNB, or a polysilazane resin (SiCN). After impregnation, a fibrous preform intended to constitute the fibrous reinforcement of the part to be produced, and having a shape substantially corresponding to that of this part, is shaped by conformation of the fibrous texture using a holding tooling .

The shaping of the fibrous preform is preferably accompanied by a compacting of the fibrous structure in order to increase the volume content of fibers in the composite material of the part to be produced.

After forming the preform, the crosslinking of the resin is performed, or completed if there has been pre-crosslinking, the preform being in a tool.

Then, the consolidation is completed by a heat treatment of pyrolysis of the resin. The pyrolysis is carried out at a temperature of, for example, about 900 ° C to 1000 ° C.

Consolidation can also be achieved by chemical vapor infiltration (CVI).

After this consolidation, the densification of the fibrous preform by a ceramic matrix is continued (step 30).

The densification is advantageously carried out by chemical vapor infiltration (CVI), the parameters of the CVI process and the nature of the reaction gas phase being adapted to the nature of the matrix to be formed. It is thus possible to chain in the same oven the pyrolysis operations of the consolidation and densification resin.

The ceramic matrix formed by CVI may be an SiC matrix, or an at least partially self-healing matrix, such as a silicon-boron-carbon matrix (Si-BC) or a boron carbide (B ₄ C) matrix, or again a sequenced matrix with alternate matrix phases in non-healing ceramics and healing ceramics. We can refer in particular to documents FR 2 401 888, US 5 246 736, US 5 965 266, US 6 068 930 and US 6 291 058.

The ceramic matrix may be deposited in several successive infiltration cycles with between each cycle a machining operation to reopen the porosity of the surface material and facilitate the deposition of the matrix in the fibrous reinforcement.

As illustrated in FIG. 3, a preform is thus obtained

300 of CMC composite material (fibrous reinforcement densified by a matrix at least partially ceramic) with a blade preform portion 320, a foot preform portion 330 (with stilt preform), a platform preform portion 340 and a blade bead preform portion 350. According to the method of the invention, the preform 300 has dimensions smaller than that of the blade 100 of FIG. 1 to be produced.

FIG. 4 shows the surface state of a portion of a CMC part made from a multilayer fibrous texture of three-dimensional weaving of SiC fibers (Guipex® base satin of 8) consolidated, shaped and densified following the method described above. As measured in FIG. 5, the piece has on the surface both corrugations of more than 200 μm in amplitude and a roughness level of the order of 5 μm. Such a surface irregularity can not possibly allow to use the room as such for aerodynamic applications. By comparison, Figure 6 shows a measurement of the surface condition of a blade of a low-pressure stage of an aircraft engine, the latter having been made of metallic material. Note that this blade has no surface ripple and has a mean level of roughness of the order of 1 pm.

For this purpose and in accordance with the process of the invention, a ceramic casing is formed around the preform made of composite material. In the example described here, the ceramic casing is made by flash sintering of a powder.

As illustrated in FIGS. 7A and 7B, the preform made of composite material 300 is placed in a sintering conductive mold having internal dimensions and shape corresponding to those of the blade to be produced with interposition in the space between the preform 300 and the inner surface of the mold of a ceramic powder (steps 40 and 50). More specifically, the composite material preform 300 is first placed in a first half 401 of the sintering mold 400 on a bed of ceramic powder 50 deposited on the inner surface 401a of the first half 401 of the mold (FIG. 7A), then is covered on its exposed face of the same powder 50 before closing the second half 402 of the mold 400 (Figure 7B). The inner surfaces 401a and 402a of the mold define, once the two halves 401 and 402 of the mold together, dimensions and a shape corresponding to those of the blade to be manufactured. The ceramic powder 50 comprises one or more constituents (mixture of powders) of the ceramic of the envelope to be produced and possibly other elements also in the form of powder such as sintering agents.

The powder 50 is then stabilized / densified by "flash sintering" or "SPS" ("Spark Plasma Sintering"). The difference between conventional hot pressing (standard sintering) and flash sintering is that the heat source is not external but that an electric current (continuous - continuous pulsed - or alternating) applied via electrodes passes through the conductive mold. The "flash sintering" consists of a heat treatment under pressure with passage of an electric current that consolidates the envelope by forming bonds between grains without total melting thereof. This solder made by diffusion of material, is accompanied by a densification, that is to say a decrease in the porosity rate and hardening which gives cohesion to the envelope shaped. In addition, the attachment of the ceramic envelope on the preform made of composite material is ensured by the chemical reactions between the oxides formed during sintering and the material of the preform.

A device for implementing this flash sintering is in particular marketed by Sumitomo Electric Industries and makes it possible to subject the powder 50 to pulses (3.3 ms) of direct electrical current (typically 0-10 V, 1-5 kA ) while applying a pressure of several tens of MPa (up to 150 MPa) and this in a range of temperatures ranging from room temperature up to 2000 ° C. Flash sintering is usually done under vacuum but it is possible to work under an inert atmosphere (nitrogen, argon).

The same sintering cycle can be used as a reference for the flash sintering densification of the various compositions of the refractory material according to the invention, only the final sintering temperature is modified as a function of the refraction of the components to be sintered.

The temperature parameters chosen for the sintering cycle are, for example: a rise at 600 ° C. in 3 minutes, followed by a rise to the sintering temperature with a speed of 100 ° C./min, then a plateau at this temperature for 5 minutes and finally a descent to 600 ° C in 30 minutes and then stop heating. During the cycle, a pressure of several ten Pa is applied gradually from the beginning of the temperature rise to 600 ° C to close the majority of the remaining pores and avoid densification heterogeneity in the envelope after sintering. Thus, from the beginning of the sintering, a generally dense material can be obtained, for which the contact between the grains is optimal.

The controlled cooling allows a relaxation of residual stresses of thermal origin and to avoid the presence of cracks and microcracks in the material.

The molds used are preferably graphite and are separated from the powder by a graphite sheet to prevent sticking.

Flash sintering reduces the sintering temperature by one to several hundreds of degrees compared to standard sintering (conventional hot pressing).

According to the invention and in order to ensure a good attachment of the ceramic envelope directly on the composite material preform, that is to say without underlay between the preform and the envelope, the ceramic powder used for forming the envelope has a particle size between 0.25 μιη and 25 pm. Indeed, a too fine particle size can cause premature reagglomeration phenomena which lead to the formation of particle size too large to obtain a ceramic envelope with a satisfactory structure and surface condition.

Moreover, with a particle size too high, the contact surface between the powder grains is too small to ensure good cohesion of the coating formed. In addition, if the size of the powder grains is too large, the thermal budget (ie temperature / time pair) necessary to perform the sintering may be too high and incompatible with the thermal stability temperature of the fibers (thermostability) of the preform made of composite material, which can lead to thermal degradation of the fibers considerably reducing their mechanical properties.

The composition of the powder intended to form the ceramic envelope around the fiber preform is also chosen so as to have a sintering temperature compatible with the thermal stability temperature of the fibers (thermostability) of the composite material preform.

In the case of flash sintering, the temperature rises and the sintering times being very short, it is possible to use powder compositions having sintering temperatures which can exceed the stability temperature of the fibers of the preform by 250.degree. .

A powder composition is also chosen which has a coefficient of thermal expansion close to that of the composite material of the preform, namely, in accordance with the invention, a powder having a coefficient of thermal expansion varying at most from 4.10 to ⁶ K ^{". 1} relative to the coefficient of thermal expansion of the composite material of the preform. In the case of flash sintering, a powder having a coefficient of thermal expansion varying at most from 1.5 × 10 ^-6 K ^-1 with respect to the thermal expansion coefficient of the composite material of the preform is preferably chosen. that is, a powder having a coefficient of thermal expansion of between 3.10 ^"6 K ^{" 1} and 6.10 ^"6 K ^{" 1} .

Table I below gives examples of powder compositions that can be used for making a ceramic shell by flash sintering, the table further indicating the flash sintering temperatures of these compositions as well as compatible SiC fiber types. with these flash sintering temperatures.

Table I

It should be noted that the powder compositions indicated in Table I also have a coefficient of thermal expansion which is relatively close to that of the CMC material of the preform.

At the end of the flash sintering, a blade is obtained such that the blade 100 shown in FIG. 1 comprises a core of composite material corresponding to the preform made of composite material 300 described above and an outer ceramic shell 150 defining the shape and the final dimensions of dawn.

FIGS. 8A and 8B show a part 200 corresponding to the blade portion of a blade and whose fibrous preform 210 has been covered by a ceramic casing 220 made by flash sintering of a Si / Mullite / BSAS powder as explained above.

In order to demonstrate the importance of the choice of the particle size of the ceramic powder used to form the envelope on the preform made of composite material, comparative tests were carried out.

The first test is illustrated in FIGS. 10 and 11 which show the result obtained after flash sintering of a layer of yttrium disilicate powder (Y2S12O7) produced at 1200 ° C. on a substrate made of CMC composite material, the powder used. to form the coating of Figure 10 having a particle size (D50) of 30 μm while the powder used to form the coating of Figure 11 has a particle size (D50) of 1 μm.

As can be seen very clearly in FIG. 10, the formed coating exhibits a continuous decohesion with the CMC substrate which results in poor adhesion of the coating to the substrate. Cracks are also present in the coating. This decohesion and these cracks are due mainly to the lack of contact surface between the grains because of the choice of a particle size too high (greater than 25 pm). In contrast, the coating formed in FIG. 10 with a powder having a particle size of 1 μm, i.e. between 0.25 μm and 25 μm, does not show any decohesion with the substrate or cracks.

The second test is illustrated in FIGS. 12 and 13 which show the result obtained after flash sintering of a layer of yttrium disilicate powder produced at 1200 ° C. on a substrate made of CMC composite material, the powder used to form the coating of Figure 10 having a particle size (D50) of 0.2 μm while the powder used to form the coating of Figure 11 has a particle size (D50) of 0.5 μm.

As can be seen very clearly in FIG. 12, the coating formed has a continuous decohesion with the CMC substrate due to premature re-agglomeration phenomena of the grains during the sintering of grains of a size that is too small (less than 0 , 25 μm). In contrast, the coating formed in FIG. 13 with a powder having a particle size of 0.5 μm, that is to say between 0.25 m and 25 μητι, does not show any decohesion with the substrate nor cracks.

According to an alternative embodiment of the method according to the invention described in FIG. 9, a composite material preform having dimensions smaller than those of the aerodynamic part to be formed in the same manner as described hereinabove is manufactured first. before, namely formation of a fibrous texture (step 110), consolidation and shaping of the fibrous texture (step 120) and densification of the preform (step 130). In this alternative embodiment, the composite material preform is placed in a mold whose inner surface corresponds to the shape and the final dimensions of the aerodynamic part to be produced (step 130), the ceramic powder intended to form the envelope ceramic is not yet present in the mold. Before it is introduced into the mold, the ceramic powder is premixed with a binder (step 140). The mixture is then injected into the mold in order to fill the space between the preform made of composite material and the internal surface of the mold (step 150). During the injection of the mixture, the piece is held in the center of the mold preferably at sacrificial portions or participating in the aerodynamics of the piece as the end portions.

Once the mixture has been injected, the binder is crosslinked (step 160). The part is then demolded and the ceramic casing formed on the preform by chemical vapor infiltration (CVI) of SiC (step 170) is stabilized, this CVI also making it possible to grip the ceramic envelope on the preform made of composite material.

As indicated above, a powder composition having a coefficient of thermal expansion varying at most from 4.10 ^"6 K ^{" 1} with respect to the thermal expansion coefficient of the composite material of the preform, that is to say between 0, is chosen. , 5.10 ^"6 K ^{" 1} and 8.5 × 10 ^" ⁶ K ^{" 1} , and a particle size of between 0.25 μητι and 25 μm. If the powder has a particle size less than 0.25 μm, it then becomes difficult for the infiltration gases of the CVI to penetrate into the powder layer and thus ensure good stabilization of the latter. In addition, if the particle size of the powder exceeds 25 μητι, there is no more surface area. sufficient contact between the grains to ensure satisfactory cohesion of the ceramic shell formed.

The binder used may in particular be a liquid organosilicon liquid precursor compound such as a polycarbosilane resin (PCS) or polysilazane (Ceraset® PSZ20), or a thermoplastic resin such as polyvinyl acetate (PVA) or polyvinyl of butyral (PVB).

According to yet another variant, a mat of ceramic fibers, for example of SiC fibers, is placed around the preform made of composite material before the injection of the ceramic-binder powder mixture into the mold in order to reinforce the mechanical strength of the envelope ceramic formed. Likewise, short ceramic fibers, for example short SiC fibers, can be introduced into the ceramic-binder powder mixture in order to reinforce the ceramic of the envelope.

The ceramic powders used to produce the ceramic powder-injected binder mixture may be chosen in particular from the powder compositions indicated in Table I above.

The ceramic of the envelope is further selected according to the conditions of use of the aerodynamic part. In particular, it must be able to withstand the operating temperatures of the room and have a service life at least equal to that defined for the room. For this purpose, a ceramic having a melting point higher than the maximum temperature of use of the part is chosen. In the case for example of parts constituting gas turbine blades, the maximum temperatures encountered by these parts can reach 1100 ° C. In this case, the ceramic of the envelope of the blade has a melting temperature greater than or equal to 1300 ° C.

For the constituents of the ceramic powder used in the invention, a particle size will be selected adapted to the final surface state of the target part. Powderes whose grain size is less than 20 μm are preferably chosen. Nanopowders may also be used because they reduce the sintering temperature.

Whatever the implementation of the process, an aerodynamic part is obtained which has very good mechanical and structural characteristics conferred by the composite material preform and a surface state comparable to that obtained with metallic materials, and this without final machining because the part has its shape and final dimensions immediately after overmoulding of the cylindrical envelope.

Furthermore, according to the envisaged conditions of use, the fibers of the fibrous reinforcement of the preform made of composite material of the part may be of a material other than a ceramic, for example carbon, and the matrix may be made of a material other than than a ceramic, for example carbon or a resin. As described above, the aerodynamic part preform according to the invention may in particular be made of ceramic matrix composite material (CMC) which is a material formed of a carbon fiber or ceramic reinforcement densified by a matrix at least partially ceramic such as carbon-carbon / silicon carbide (CC / SiC), carbon-silicon carbide (C / SiC), silicon carbide-silicon carbide (SiC / SiC). The preform may also be made of carbon-carbon (C / C) composite material which, in known manner, is a material formed of a carbon fiber reinforcement densified by a carbon matrix. The invention is of course also applicable to the manufacture of blades made of organic matrix composite material (CMO) such as that obtained, for example, from a high performance epoxy resin.

Claims

A method of manufacturing an aerodynamic workpiece comprising:

- The interposition of a ceramic powder between the composite material preform and the inner surface of the mold, said ceramic powder having a coefficient of thermal expansion varying at most 4.10 ^{~ 6} .K ^{~ 1} relative to the coefficient of thermal expansion of the composite material of the preform, said ceramic powder further having a particle size of between 0.25 μm and 25 μm,

stabilizing said powder so as to form around the composite material preform a ceramic envelope with the dimensions and the shape of the aerodynamic part to be produced.

2. Method according to claim 1, characterized in that the stabilization of the ceramic powder is carried out by flash sintering.

3. Method according to claim 2, characterized in that one chooses a ceramic powder composition having a maximum flash sintering temperature greater than 250 ° C to the temperature of thermal stability of the fibers of the preform of composite material.

4. Method according to claim 1, characterized in that the interposition of the ceramic powder is obtained by producing a mixture of said ceramic powder with a binder and injecting the mixture between the preform and the inner surface of the mold and in that the stabilization of said powder is performed by chemical vapor infiltration of silicon carbide.

5. Method according to claim 4, characterized in that the binder is selected from at least: an organosilicon compound and a fugitive resin.

6. Method according to claims 4 and 5, characterized in that, before the injection of the mixture of binder and ceramic powder, there is a mat of ceramic fibers around the fiber preform.

7. Method according to claims 4 and 5, characterized in that the mixture of binder and ceramic powder further comprises ceramic fibers.

8. Method according to any one of claims 1 to 7, characterized in that the ceramic powder comprises nanopowders.

9. Method according to any one of claims 1 to 8, characterized in that the aerodynamic part is a fixed blade or mobile turbine engine.

10. Method according to any one of claims 1 to 8, characterized in that the preform is made of ceramic matrix composite material (CMC).