Fabrication of composite articles having an infiltrated matrix

A composite article is fabricated by preparing a porous first region piece having a first reinforcement material with no matrix, and rigidizing the first reinforcement material to form a porous first coated preform. At least one second-region layer of a second reinforcement material having no matrix is applied in contact with at least a portion of the first coated preform. The second reinforcement material is rigidized, to form a second coated preform. A second-region matrix material is deposited into at least some of the porosity of the second coated preform. The reinforcement material may be silicon carbide fibers, and the infiltrated matrix may be silicon carbide.

BACKGROUND OF THE INVENTION
 This invention relates to the fabrication of composite articles, and, more
 particularly, to the fabrication of such articles using a fibrous ceramic
 reinforcing material that is infiltrated with a ceramic matrix material.
 In an aircraft gas turbine (jet) engine, air is drawn into the front of the
 engine, compressed by a shaft-mounted compressor, and mixed with fuel. The
 mixture is combusted, and the resulting hot exhaust gases are passed
 through a turbine mounted on the same shaft. The flow of gas turns the
 turbine, which turns the shaft and provides power to the compressor. The
 hot exhaust gases flow from the back of the engine, driving it and the
 aircraft forwardly.
 The hotter the exhaust gases, the more efficient is the operation of the
 jet engine. There is thus an incentive to raise the exhaust gas
 temperature. However, the maximum temperature of the exhaust gases is
 normally limited by the materials used to fabricate the turbine vanes and
 turbine blades of the turbine. In current engines, the turbine vanes and
 blades are made of nickel-based superalloys and can operate at
 temperatures of up to 1900-2100.degree. F.
 Many approaches have been used to increase the operating temperature limit
 of the turbine blades and vanes. The compositions and processing of the
 materials themselves have been improved. Physical cooling techniques are
 used. In one widely used approach, internal cooling channels are provided
 within the components, and cool air is forced through the channels during
 engine operation.
 In another approach, ceramic or ceramic composite materials have been used
 to fabricate some of the hot section components. Most ceramics have very
 limited fracture toughness, and therefore ceramic composite materials have
 been considered for such structures. A ceramic composite material of
 current interest is silicon carbide fibers embedded in a silicon carbide
 matrix. In one approach, articles are fabricated by collating silicon
 carbide fibers on a tool, rigidizing the silicon carbide fibers to form a
 coated preform, and then producing a silicon carbide matrix in the coated
 preform by chemical vapor deposition (i.e., chemical vapor infiltration)
 or melt infiltration.
 While operable for the fabrication of many articles, the present inventors
 have recognized that the current manufacturing process has shortcomings
 when used in the fabrication of other articles. For example, if the
 article is quite thick, the production of the matrix is slow or may not be
 possible. Some hollow articles and articles containing cooling channels,
 such as turbine blades or vanes, cannot be readily prepared by the
 conventional procedure.
 There is a need for an improved approach to the fabrication of composite
 articles to allow greater flexibility in the preparation of complex and
 thick sections. The present invention fulfills this need, and further
 provides related advantages.
 SUMMARY OF THE INVENTION
 The present invention provides a method for the fabrication of composite
 articles, and articles prepared by the method. This approach allows the
 fabrication of significantly thicker articles than possible with
 conventional procedures. Articles with corners that require doublers at
 the interior side of the comer may be prepared. Hollow articles with
 multiple internal cooling passages and externally connecting orifices may
 be fabricated more easily than with the conventional approach. The
 finished articles are sound and of high quality throughout.
 In accordance with the invention, a method for fabricating a composite
 article comprises the steps of preparing a porous first region piece
 comprising a first reinforcement material having no matrix, and thereafter
 rigidizing the first reinforcement material to form a porous first coated
 preform. The method further includes applying at least one second-region
 layer of a second reinforcement material having no matrix in contact with
 at least a portion of the first coated preform, and thereafter rigidizing
 the second reinforcement material, to form a second coated preform. A
 second-region matrix material is deposited into at least some of the
 porosity of the second coated preform. The reinforcement material is
 preferably silicon carbide fiber, and the matrix material preferably
 comprises silicon carbide, either polycrystalline silicon carbide, or a
 mixture of polycrystalline silicon carbide and silicon.
 The present approach may be used to fabricate solid (i.e., not hollow)
 articles or hollow articles such as turbine components with internal
 cooling passages and externally connecting cooling orifices. In one
 embodiment of this approach, the first coated preform is hollow, defining
 one internal passage, and/or the second-region layer of the second
 reinforcement material may contact some, but not all, of the outer
 periphery of the first coated preform. In another embodiment, to make a
 hollow article with an internal rib defining a second large internal
 passage, a third coated preform is made and assembled with the first
 coated preform. Both the first and third coated preforms are overlaid with
 the second reinforcement material, and the matrix material is infiltrated
 to form the shell of the hollow article. With the prior approach, it is
 quite difficult to fabricate an article having an internal rib.
 Optionally, a first-region matrix material may be deposited into at least
 some of the porosity of the first coated preform after the step of
 rigidizing the first reinforcement material and before the step of
 applying at least one second-region layer. This allows a progressive
 infiltration of the matrix material, or the use of a different matrix
 material. The thickness of some articles made by the prior approach is
 limited by the ability to infiltrate the matrix material into the porosity
 of the fibrous coated preform, and the present approach overcomes that
 limitation so that thicker articles may be made than heretofore possible.
 Other features and advantages of the present invention will be apparent
 from the following more detailed description of the preferred embodiment,
 taken in conjunction with the accompanying drawings, which illustrate, by
 way of example, the principles of the invention. The scope of the
 invention is not, however, limited to this preferred embodiment.

DETAILED DESCRIPTION OF THE INVENTION
 FIG. 1 depicts a composite article 20 in the form of two panels joined at a
 relatively sharp corner, numeral 22, with a reinforcing doubler 24 on the
 interior side of the sharp corner. This article 20 may be difficult or
 impossible to fabricate from some materials, such as a composite material
 of silicon carbide (SiC) fibers within an infiltrated polycrystalline
 silicon carbide matrix. The difficulty may arise both from the geometry of
 the positioning of the doubler 24 on the interior side of the sharp
 corner, and from the difficulty in infiltrating silicon carbide matrix
 material through the total thickness of the article 20 in the region of
 the doubler 24.
 FIG. 2 is a block diagram of an approach for fabricating articles by the
 approach of the invention, such as the article 20 of FIG. 1, and other
 articles as will be discussed subsequently. The method will be described
 here in relation to the article 20, but is not so limited. A first-region
 piece is first prepared, numeral 30. In this case, the first-region piece
 has the shape of the panels and corner 22. It is formed of a first
 reinforcement material with no matrix material present. The reinforcement
 material is preferably ceramic fibers, most preferably silicon carbide
 fibers. The silicon carbide fibers may be unidirectional or cross-plied
 two dimensionally, in one or more layers, or even furnished in a
 three-dimensional arrangement. In one embodiment, the silicon carbide
 fibers are furnished as a single two-dimensional layer of woven cloth,
 termed a "ply". Typically, several plies are collated together to provide
 the desired thickness of the first-region piece, usually on appropriate
 tooling that defines the shape of the first-region piece.
 The first reinforcement piece is rigidized, numeral 32. In the initial
 first-region piece prepared in step 30, the reinforcement fibers are free
 to move relative to each other by small amounts, which permits the plies
 of the first-region piece to be collated but prevents easy handling of the
 first-region piece. In the rigidizing step 32, a bonding material is
 applied to the reinforcement fibers at their points of contact to
 transform the first-region piece into a relatively rigid structure. As
 shown in FIG. 3, in the preferred approach the reinforcing fibers 50 are
 present in a weave with porosity 52 therebetween at this stage. At a later
 stage, the porosity is filled with a matrix material. In the rigidizing
 step 32, the bonding material 54 is applied to the reinforcing fibers 50,
 so that, at their points of contact 56 with neighboring reinforcing
 fibers, they are bound together. In a preferred rigidizing process, a thin
 layer of boron nitride is first deposited into the porous first-region
 piece, so that it coats the reinforcing fibers 50, both generally and at
 the points of contact 56. Silicon carbide is thereafter deposited
 overlying the boron nitride, both generally and at the points of contact
 56. The deposited boron nitride and silicon carbide are the bonding
 material 54 at the points of contact 56 of the reinforcing fibers 50. The
 result, as depicted in FIG. 3, is a bonded structure, known as a first
 coated preform, that has sufficient rigidity to be handled in subsequent
 processing, although its mechanical properties are not sufficient for
 service applications. The deposition of the boron nitride and silicon
 carbide layers are preferably accomplished by chemical vapor deposition
 (CVD), which, in this application, is often termed chemical vapor
 infiltration (CVI). This deposition technique is well known, see, for
 example, D. P. Stinton et al., "Fabrication of Ceramic-Ceramic Composites
 by Chemical Vapor Deposition", Ceramic Engineering Science Proceedings,
 Vol. 5 {7-8}, pages 668-676 (1984).
 Optionally, a first-region matrix material is deposited into the porosity
 52 of the first coated preform, numeral 34. For many articles the matrix
 material may be deposited in a single step, to be described subsequently
 as step 40. That approach is preferred, because for some articles it is
 less time consuming to perform a single deposition step than two
 deposition steps. However, in some cases the final structure is too thick
 or complex to permit deposition of the matrix for the entire article in
 one step. The practical value of the thickness limit of the part for
 successful deposition of the matrix depends on several factors, such as
 the arrangement of the fiber reinforcement, the volume fraction of the
 fiber reinforcement (i.e., the amount of open space between the fibers
 during deposition of the matrix), the geometry of the part, the required
 composition of the matrix, and the method used to deposit the matrix.
 Therefore, no single firm value may be stated, but upper thickness limits
 are observed for matrix deposition for each particular combination of
 variables. For a final part having a high fiber volume fraction of about
 40 percent or cross-woven fibers, the practical maximum part thickness for
 chemical vapor infiltration is about 1/8 inch and the practical maximum
 part thickness for melt infiltration is about 1/4-1 inch.
 A virtue of the present sequential fabrication process is that it permits
 components to be infiltrated with the matrix material at intermediate
 stages, in this case in optional step 34, so that thicker parts may be
 prepared and also parts of configurations not possible with the
 conventional approach. Another advantage is that the matrix material
 introduced into the first coated preform may be different than that used
 in other portions of the structure, although in most cases the same matrix
 material is used throughout. The procedure used in this optional step 34
 is the same as that used in step 40, which subsequent description is
 incorporated here.
 At least one second-region layer of a second reinforcement material is
 applied to the first coated preform (or infiltrated first coated preform
 if step 34 is employed), numeral 36. The second reinforcement material is
 preferably the same type of material as the first reinforcement material,
 although the present approach permits the use of a different second
 reinforcement material. In the preferred case, the second reinforcement
 fiber material is provided as a plies of a two-dimensional weave of
 reinforcement fibers, most preferably silicon carbide fibers. The plies
 are applied to the surface of the first coated preform (or infiltrated
 first coated preform if step 34 is used), so that they contact at least a
 portion of the surface of the first coated preform. In the article of FIG.
 1, the plies of the second reinforcement fibers are applied to the inner
 surface of the panels and corner 22, near to the inner surface of the
 corner material, in the region that will eventually be the reinforcing
 doubler 24.
 The newly applied second reinforcement material is rigidized, numeral 38.
 The rigidizing step 38 is performed by the same approach as the rigidizing
 step 32, which discussion is incorporated here. The resulting porous
 structure is termed a second coated preform. It has a microstructure like
 that of FIG. 3, which description is incorporated here. The second coated
 preform includes the reinforcement material applied in step 36, and also
 the reinforcement material applied earlier in the process but not
 previously infiltrated with matrix material.
 A second-region matrix material is deposited into at least some of the
 porosity of the second coated preform, numeral 40. The second-region
 matrix material is preferably silicon carbide, deposited in a
 polycrystalline form into the porosity, or a mixture of silicon and
 polycrystalline silicon carbide. Deposition of the matrix material may be
 accomplished by any operable approach that allows the matrix material to
 be infiltrated into the porosity of the coated preform. One technique is
 to deposit the matrix material by chemical vapor deposition, using
 essentially the same approach as in the rigidizing steps 32 and 38. In
 another approach, the matrix material is deposited by melt infiltration.
 In one method, a slurry of carbon, carbon-containing resin, or other
 carbonaceous material, and, optionally, silicon carbide particulate, is
 introduced into the porosity, and molten silicon is thereafter infiltrated
 into the remaining space to react with the carbonaceous material to form
 silicon carbide. The amount of silicon may be stoichiometric, so that the
 matrix is silicon carbide. An excess of silicon may instead be used, so
 that the final structure is a mixture of reacted silicon carbide and
 unreacted silicon. In a typical case of the latter embodiment, the silicon
 carbide:silicon weight ratio ranges from about 95:5 to 50:50. Examples of
 deposition processes, often generally termed Silcomp processes, are
 disclosed in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350, whose
 disclosures are incorporated by reference.
 At least some of the matrix material of the first region may be deposited
 in the optional step 34. It is preferred, however, that at least some of
 the matrix material of the first region also be deposited during the
 deposition step 40, so as to provide a bonding between the first region
 and the second region.
 The resulting article 20 comprises a composite material structure, with
 silicon carbide fibers embedded in a silicon carbide matrix in the
 preferred embodiment. The article may be made thicker than is possible
 with conventional fabrication technology, as discussed above.
 The present approach also permits the fabrication of complex composite
 articles which are difficult or impossible to fabricate by conventional
 fabrication technology. FIG. 4 depicts a component article of a gas
 turbine engine such as a turbine blade or turbine vane, and in this
 illustration a hollow composite low pressure turbine vane 60 that is to be
 fabricated. The turbine vane 60 includes an airfoil 62 against which the
 flow of hot exhaust gas is directed. The turbine vane 60 is hollow for
 reduced weight and to provide a channel for a flow of cooling air to pass
 through the interior of the turbine vane and exit through orifices 64 in
 the surface of the airfoil 62. In operation, a flow of cooling air is
 directed through the cooling channel and out the orifices 64, to reduce
 the temperature of the airfoil 22.
 FIG. 5 is a sectional view of the turbine vane 60 of FIG. 4. The turbine
 vane 60 is hollow, with three channels 66a, 66b, and 66c through which
 cooling air flows axially through the turbine vane 60 to exit at the
 orifices 64. The use of three channels 66a, 66b, and 66c is desirable in
 order to achieve the proper pressure distribution and flow through the
 various orifices 64. To define the three channels 66a, 66b, and 66c, the
 hollow interior of the turbine vane is divided by two through-thickness
 ribs 68a and 68b extending between the outer walls 70 of the turbine vane
 60. The outer walls 70 of the turbine vane 60 may be fabricated as a
 silicon carbide fiber/silicon carbide matrix composite material using
 conventional technology. However, it is difficult or impossible to form
 the ribs 68 using the conventional approach, because the rib portion of
 the coated preform cannot be readily produced, and because the rib-portion
 of the coated preform could not be readily infiltrated with the matrix
 material in any event.
 Such a complex article as the turbine vane 60 may be fabricated using the
 present approach, following the procedure generally illustrated in FIG. 2,
 whose description is incorporated here. An insert 80, depicted in FIG. 6,
 is prepared as the first region piece in step 30. The insert 80 is
 prepared by collating first reinforcement material onto a piece of tooling
 in the shape of a mandrel 82. In the preferred case, plies of silicon
 carbide fabric are collated onto the mandrel. The mandrel 82 is selected
 so that the insert 80 has an inner surface 84 whose shape and size defines
 an inner surface 86 of a portion of the final article, specifically the
 channel 66b of FIG. 5. While on the mandrel, the first reinforcement
 material is rigidized, numeral 32, yielding the coated preform. The porous
 coated preform has sufficient strength and rigidity to be processed at
 this stage, as by machining openings 88 that are required to communicate
 the air flow to the orifices 64 in the final article.
 After the coated preform is complete, the mandrel 82 is removed by a
 physical, chemical, or thermal process. For example, the mandrel 82 may be
 machined out of the interior of the insert 80, or may be dissolved, etched
 away, or vaporized, depending upon its material of construction.
 Optionally, matrix material may be deposited into the porosity of the
 insert coated preform, numeral 34.
 The second reinforcement material is applied overlying the insert coated
 preform, step 36. In the preferred case, the second reinforcement material
 is additional plies of silicon carbide fabric, as described previously.
 The total thickness of the article at any location is the sum of the
 thickness of the insert and the overlying material. The total thickness of
 material required by strength considerations is allocated between the
 insert and the overlying material in the design process, with typically
 about half the total number of plies in the insert 80 and half in the
 overlying material.
 The second reinforcement material may be applied so as to contact all of
 the insert 80, or so as to contact only a portion of the surface of the
 insert 80. By applying the second reinforcement material so as to contact
 only a portion of the surface of the insert, and bowing the second
 reinforcement material outwardly to define the surface shape of the
 airfoil 62, the hollow interior of the article may be subdivided to define
 the channels 66a, 66b, and 66c. The ends of the insert 60 thereby define
 the ribs 68a and 68b, as shown in FIG. 7.
 The second reinforcement material is rigidized, step 38. The resulting
 second coated preform, shown in FIG. 7, includes the coated preform region
 of the insert 80, and a coated preform region of a shell 90 of the
 rigidized second reinforcement material overlying the insert 80. The shape
 and thickness to which the second reinforcement material was applied in
 step 36 defines the shape and size of the airfoil 62 of the turbine vane
 60.
 Matrix material is deposited into the porosity of the second reinforcement
 material and, if step 34 was not performed or performed so as to fill only
 a portion of the porosity of the first coated preform, into the porosity
 of the first reinforcement material, step 40. The preferred silicon
 carbide or silicon/silicon carbide matrix material is deposited as
 described previously.
 The fabrication of the turbine vane 60 is completed by final machining,
 such as the drilling of the orifices 64 aligned with the openings 88 or
 fine machining the exterior surface of the shell 90 so as to achieve a
 precise size and shape for the airfoil. The final machining may be
 performed either before or after the deposition step 40, as may be
 appropriate for particular final machining operations. The final
 fabrication may also include removing any stray matrix material and
 general cleaning.
 FIG. 9 illustrates an even more complex turbine vane 100, wherein the
 hollow interior is divided into four channels rather than three channels,
 and including an additional transverse rib, to achieve a different
 distribution of the cooling air flow and more structural strength in the
 final turbine vane. The fabrication of this turbine vane 100 is similar to
 that of the turbine vane 60, and the prior discussion is incorporated
 here. In this case, however, two inserts 80a and 80b are prepared and
 positioned together so as to form a central transverse rib 102 that
 further subdivides the hollow interior of the turbine vane 100 and
 provides additional structural strength.
 An airfoil section has been prepared using the approach of the invention.
 The microstructure of this section was good, with a fully dense matrix and
 the various coatings and layers uniform. Optional step 34 was not employed
 in this prototype.
 Although a particular embodiment of the invention has been described in
 detail for purposes of illustration, various modifications and
 enhancements may be made without departing from the spirit and scope of
 the invention. Accordingly, the invention is not to be limited except as
 by the appended claims.