Ceramic matrix composite blade with integral platform structures and method and tool for its manufacture

Abstract

A method of manufacturing a ceramic matrix composite blade (10), comprising: Placing at least one layer of a plurality of fiber ceramic layers in a preselected arrangement on a tool support surface (206) to form a blade and dovetail preform; Laying at least one further layer of the plurality of fiber ceramic layers onto the airfoil and dovetail preform in a second preselected arrangement to form an integral platform preform, wherein a plurality of the further fiber ceramic layers are stacked to form an unstiffened blade preform together with the blade and dovetail preform; Stiffening the unstiffened blade preform to form a stiffened blade preform, wherein the airfoil (12), the dovetail (14), and the integral platform (16) comprise a substantially continuous co-stiffened matrix phase; and Compacting the stiffened blade preform to form the ceramic matrix composite blade (10) with an integral platform structure.

Description

DECLARATION REGARDING GOVERNMENTALLY SPONSORED RESEARCH OR DEVELOPMENT

[0001] Embodiments of the present disclosure were made with Government support under Contract No. N00421-05-C-0053. The Government may have certain rights. AREA

[0002] The present disclosure is directed to composite rotor blades with integral platform structures. More specifically, the present disclosure relates to ceramic matrix composite rotor blades and methods and tools for manufacturing blades with integral platform structures. BACKGROUND OF REVELATION

[0003] To increase the efficiency and performance of gas turbine engines, such as enabling increased thrust-to-weight ratios, lower emissions, and improved specific fuel consumption, engine turbines are required to operate at higher temperatures. As these higher temperatures approach or exceed the limits of the materials used to construct components in the hot section of the engine, and particularly the turbine section, new materials are desired.

[0004] As engine operating temperatures have increased, new methods have been developed for cooling high-temperature alloys, including those in combustion chambers and turbine blades. For example, ceramic thermal barrier coatings (TBCs) have been applied to the surfaces of components exposed to the hot exhaust combustion gas stream to reduce heat transfer rates, provide thermal protection to the underlying metal, and enable the component to withstand higher temperatures. These improvements helped reduce peak temperatures and thermal gradients. Cooling holes have also been added to enable film cooling to improve thermal performance or thermal protection. Ceramic matrix composites (CMCs) have also been developed as replacements for high-temperature alloys.In many cases, CMCs offered improved temperature performance and a density advantage over metals, making them the material of choice when higher operating temperatures were desired.

[0005] In the past, a number of techniques have been used to manufacture turbine engine components, such as turbine blades, using CMC. For example, silicon CMC can be formed from a fiber material infiltrated with molten silicon. Such a process is typically referred to as the Silcomp process. In this type of process, the fibers generally have diameters of about 140 µm or larger, which prevents elaborate, complex shapes, such as turbine blade components, from being manufactured by the Silcomp process.

[0006] Another approach to manufacturing CMC turbine blades is the process known as the slurry cast melt infiltration (MI) process. In a manufacturing process using the slurry cast MI process, CMCs are produced by creating layers of matched two-dimensional (2D) woven fabric with silicon carbide (SiC)-containing fibers having two weave directions at angles of substantially 90° to each other, with substantially the same number of fibers running in both directions of the fabric. The term "silicon carbide-containing fiber" refers to a fiber having a composition that includes silicon carbide, and preferably is essentially silicon carbide. For example, the fiber may include a silicon carbide core surrounded by carbon, or conversely, the fiber may have a carbon core surrounded by or encased in silicon carbide.

[0007] Other techniques for manufacturing CMC components include polymer infiltration and pyrolysis (PIP). In this process, silicon carbide fiber preforms are infiltrated with a preceramic polymer, such as polysilazane, and then heat-treated to form a SiC matrix.

[0008] Yet another method for forming CMC components may involve an oxide / oxide process. In this type of processing, aluminum or aluminosilicate fibers can be pre-impregnated and then laminated into a pre-selected geometry.

[0009] Components can also be manufactured from a carbon fiber reinforced silicon carbide matrix (C / SiC) CMC. C / SiC processing involves a carbon fiber preform constructed in a preselected geometry. When used in the SiC / SiC slurry casting process, the tool is made of a graphite material. The fiber preform is held by the tool during a chemical vapor infiltration process at approximately 1200°C, forming the C / SiC CMC component.

[0010] Current CMC blade formation processes fail to enable the formation of an integral platform. Subsequent platform formation and / or the incorporation of metallic platform structures do not provide the desired performance characteristics for the blade and may result in separation of the platform structure from the airfoil and loss of adequate sealing.

[0011] EP 1 801 354 A2 describes a composite rotor blade assembly comprising an airfoil, a base, and an integral platform, wherein the airfoil and the base are one-piece, co-directionally extending, and comprise a plurality of layered plies of a fiber composite material, wherein the plurality of layered plies of fiber composite material are aligned in the airfoil in a first selected principal direction. The platform comprises a platform base surrounding the airfoil, wherein the platform comprises a plurality of layered plies of fiber composite material penetrated by the airfoil.For production, a pre-created airfoil and base preform is inserted through an airfoil-shaped opening of a separately prefabricated platform. Joints between the airfoil and base preform and the platform preform are bonded with a compatible binder, and the resulting assembly is partially cured to form an airfoil preform. The airfoil preform is then penetrated with a melt binder and fully cured.A tool for producing a rotor blade from a ceramic matrix composite material, has support surfaces for laying fiber ceramic layers to form a platform preform, and an opening in the support surfaces for receiving an airfoil preform of the rotor blade in such a way that the airfoil preform penetrates the platform preform to be bonded together with a binder, receive a fusible binder and be co-cured.

[0012] What is needed is a composite material with an integral platform structure that is easily formed and provides the desired performance characteristics of a CMC blade. SUMMARY OF REVELATION

[0013] According to one aspect of the invention, a method for manufacturing a ceramic matrix composite blade includes laying at least one layer of a plurality of fiber ceramic layers in a preselected arrangement on a tool support surface to form an airfoil and dovetail preform, laying at least one further layer of the plurality of fiber ceramic layers on the airfoil and dovetail preform in a second preselected arrangement to form an integral platform preform, wherein a plurality of the further fiber ceramic layers are stacked to form an unstiffened blade preform together with the airfoil and dovetail preform, stiffening the unstiffened blade preform to form a stiffened blade preform, wherein the airfoil,the dovetail and the integral platform have a substantially continuous co-stiffened matrix phase, and densifying the stiffened blade preform to form a ceramic matrix composite blade with an integral platform structure.

[0014] According to another aspect of the invention, a tool for producing a ceramic matrix composite blade includes a first tool component having a first support surface adapted to receive at least one ceramic fiber support to form an airfoil and dovetail preform, and a second tool component having a second support surface adapted to receive at least one further ceramic fiber support to form an integral platform preform. In the exemplary embodiment, the first support surface is selectively configurable to permit the integral platform preform to be placed over the airfoil and dovetail preform to form a blade preform, the first support surface having a dovetail shape that is removable to alter the first support surface.

[0015] According to yet another aspect of the invention, a ceramic matrix composite blade assembly includes an airfoil, a dovetail, and an integral fiber-reinforced platform in a preselected arrangement within a ceramic matrix. The integral platform includes platform surfaces and dovetail surfaces enclosing the integral platform. The airfoil, dovetail, and integral platform all comprise a substantially continuous co-stiffened composite phase.

[0016] An advantage of the present disclosure is the improved resistance to the local stresses of gas turbine operation.

[0017] Another advantage of the present disclosure is a single tool for producing an airfoil, a dovetail, and an integral platform.

[0018] Another advantage of the present disclosure is a blade, dovetail and platform that are unitarily formed and provide improved strength.

[0019] Another advantage of the present disclosure is that conventional angel wing flow path sealing designs can be used.

[0020] Further features and advantages of the present disclosure will become apparent from the following more detailed description of the preferred embodiments, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the disclosure. BRIEF DESCRIPTION OF THE DRAWINGS Fig. 1 shows a perspective view of an exemplary embodiment of a rotor blade. Fig. 2 shows an exploded perspective view of an exemplary embodiment of a blade preform. Fig. 3 shows a perspective view of an exemplary embodiment of a rotor blade. Fig. 4 shows a perspective view of another exemplary embodiment of a blade. Fig. Figure 5 shows a perspective view of an exemplary embodiment of blade assemblies. Fig. Figure 6 shows a perspective view of an exemplary embodiment of blade assemblies. Fig. 7 shows a perspective view of an exemplary embodiment of an unassembled tool for manufacturing a blade assembly. Fig. Figure 8 shows a perspective view of an exemplary embodiment of a composite tool for manufacturing a blade assembly. Fig. 9 shows an exemplary embodiment of a method for manufacturing a rotor blade assembly.

[0021] Wherever possible, the same reference numerals are used in the drawings to identify the same elements. DETAILED DESCRIPTION OF THE INVENTION

[0022] The present disclosure relates to an article, method, and tool for manufacturing a blade including a matrix composite material.

[0023] With reference to Fig. 1: An exemplary embodiment of a blade 10 is shown, depicted as a low-pressure turbine (LPT) aircraft engine blade configured for use in a blade assembly 100. When used herein, the term "blade" encompasses an article including at least one section having an airfoil portion, a dovetail portion, and a platform, and may be a preform, a stiffened component, and / or a densified component. In other embodiments, the blade 10 is configured for use in a fan blade assembly, a compressor blade assembly, or any other suitable application. In some embodiments, the assembly may include one or more blades 10 (see Fig. 1 to 4). In alternative embodiments, the arrangement may include one or more rotor blades 10 or no rotor blades 10 (see Fig. 5 and Fig. 6). The blade assembly 100 and / or the blade 10 may be made of a ceramic matrix composite (CMC). In other embodiments, the blade assembly 100 and / or the blade 10 may include other matrix materials, such as epoxy materials (e.g., for fans), polymer composites (e.g., for compressors), or any other suitable matrix material.

[0024] As in Fig. 1, the rotor blade 10 is formed from a ceramic composite material, such as, but not limited to, silicon carbide fibers embedded in a silicon carbide matrix (SiC / SiC). The rotor blade 10 includes an airfoil 12 against which the flow of hot exhaust gas is directed. The rotor blade 10 may be attached to a turbine disk (not shown) by means of a dovetail 14 extending downwardly from the airfoil 12 and engaging a groove (not shown) of a mating geometry on the turbine disk. The dovetail 14 and the surface of the airfoil 12 form a concave intersection. In other embodiments, the rotor blade 10 may have a geometry to suit other suitable applications.

[0025] In the in the Fig. 1 to 4, the blade 10 includes integral platforms 16 on a pressure side 18 and / or a suction side 20 of a dovetail 14. This arrangement may further integrate the dovetail 14 with the airfoil 12 and the dovetail 14 by incorporating fiber reinforcement in a preselected arrangement within a ceramic matrix, thereby forming a substantially continuous, co-stiffened matrix phase, as described below. With reference to Fig. 2: The blade 10 may have an inner layer, e.g., the airfoil 12, formed by laid-up plies in a preselected arrangement in a uniform arrangement with a dovetail 14 formed by laid-up plies in a second preselected arrangement. The matrix material 28 may be laid over the airfoil 12 and selectively or partially over the dovetail 14. The integral platform 16 may be formed in a third preselected arrangement by a third set of plies, e.g., integral platform plies 22. The plies are layered to form the composite preform. The ply material that may be used includes a pre-impregnated material consisting of ceramic fibers or woven or braided ceramic fiber ply material, or layered ceramic fiber ropes.The third preselected arrangement may be configured to be overlaid on the dovetail 14, as explained below. The platform surfaces 24 and the dovetail surfaces 26 may be overlaid on integral platform sheets 22, thereby enclosing the integral platform 16.

[0026] With reference to Fig. 3: The rotor blade 10 may have pins 30. The pins 30 may be pre-impregnated SiC / SiC pins in age-hardened form and / or SiC / SiC-CMC pins. Alternatively, the pins 30 may contain carbon or graphite materials, e.g., in C / SiC-CMC rotor blades 10. Furthermore, the pins 30 may contain a pre-impregnated oxide / oxide or a solidified oxide, e.g., in oxide / oxide-CMC rotor blades 10. The pins 30 may be arranged below the integral platform 16 or on a platform surface 24. With reference to Fig. 4: The rotor blade 10 may have angel wings 32. The pins 30 provide fiber reinforcement through the thickness of the platform 16 and the dovetail 14.

[0027] With reference to the Fig. 5 and Fig. 6: The blade assembly 100 may include an integral platform 16 after manufacture. The term "blade replacement element" encompasses an article having a geometry suitable for forming a blade opening 36, and may be made of any suitable material, including, but not limited to, metal. With reference to Fig. 5: The platform 16 may extend around the blade opening 36 formed during manufacture by the blade 10 or the blade replacement element. In this embodiment, the platform 16 serves as the platform for the suction side 20 and the pressure side 18 of the blade 10. The platform 16 may comprise a pre-impregnated SiC / SiC CMC material or, as described above, any other suitable material. In yet another embodiment, shown in Fig. 6 can be used in Fig. 5 shown embodiment angel wings 32 can be added, which Fig. 4 are similar.

[0028] With reference to the Fig. 7 and Fig. 8: Exemplary embodiments of a tool for manufacturing the rotor blade 10 are shown. The tool 200 may be used to manufacture the rotor blade 10 or the rotor blade assembly 100, which include an airfoil and dovetail preform and an integral platform preform. In general, the tool 200 may include a first component and a second component. The first tool component may include a first support surface 206 adapted to receive at least one ceramic fiber layer to form the airfoil and dovetail preform. The second tool component may include a second support surface 208 adapted to receive at least one further ceramic fiber layer to form the integral platform preform. The first support surface 206 and the second support surface 208 are not limited to the Fig. 7 and Fig. 8, but may also include other surfaces, including surfaces of the opposing sides 202, 204 or other components or inserts used with the tool 200. In one embodiment, the first support surface 206 may be selectively configurable to facilitate the placement of the integral platform preform onto the airfoil and dovetail preform to form a preform of the blade 10.

[0029] An exemplary tool 200 as used in the Fig. 7 and Fig. 8, includes a first group of opposed side pieces 202, 204 configured to abut and tighten against one another. As shown, the side pieces 202, 204 may be arranged as a mold for the blade 10 or a portion for supporting the blade replacement element. The side pieces 202, 204 may have a first bearing surface 206 configured to facilitate the production of the desired shape for the blade 10. The tool 200 further includes a second group of opposed side pieces 208, 210 configured to exert pressure on the airfoil 12 or the dovetail 14 (or, in alternative embodiments, on the blade replacement element).The tool 200 may include a dovetail mold 212 and / or a bridge 214 or other structures to provide a selectively configurable surface for applying preform material, such as a ceramic fiber material. In one embodiment, the dovetail mold 212 may further form a support surface, e.g., the first support surface. In another embodiment, the dovetail mold 212 is configured to co-stiffen the airfoil and dovetail preform and the integral platform preform.

[0030] With reference to Fig. 9: An exemplary method for manufacturing the rotor blade 10 using the tool 20 is shown (method 300). First, a tool (e.g., the tool 200) is prepared (step 302). The preparation of the tool may include, for example, cleaning the tool with acetone and / or applying a release agent to the dovetail shape 212. As can be appreciated, the release agent may also be applied elsewhere on the tool 200.

[0031] With reference to Fig. 9: Next, the placement of the blade 12 and the dovetail 14 on the first support surface 206 (see e.g. the Fig. 7 and Fig. 8) are performed (step 304). Laying up may involve at least one of a number of fiber ceramic layers in a preselected arrangement forming an airfoil and dovetail preform. Laying up an airfoil 12 may involve laying up ceramic fiber layers in a manner known in the art for applying ceramic fiber layers to a non-variable surface. Layers may be added or removed as needed to achieve the desired thickness.

[0032] Thereafter, the pressure side 18 of the platform 16 is laid up, for example, onto the support surface 208 (step 306). Laying up may involve at least one further layer of a number of fiber ceramic layers in a second preselected arrangement forming an integral platform preform to form an unstiffened blade preform. Laying up the platform 16 of the pressure side 18 may involve, for example, abutting the fiber layers and matrix layers against the bridge 214, removing additional matrix material, and adding or removing layers as needed to achieve the desired thickness. In an alternative embodiment for using the tool 200 to manufacture the blade assembly 100, a blade replacement element is located in the tool 200, and the layup of the airfoil 12 and dovetail 14 is omitted.In the alternative embodiment, the platform 16 is applied to the blade replacement element, for example, with a release material on the blade replacement element.

[0033] Again with reference to Fig. 9: Next, a mold (e.g., dovetail mold 212) may be removed (step 314), allowing the deposition of the platform 16 of the suction side 20 (step 318). Deposition of the platform 16 of the suction side 20 may include, for example, deposition of fiber layers and matrix layers on surfaces of the tool 200, including surfaces of the tool side 204. In other embodiments, additional structures or components may be used to form the surface onto which the platform 16 of the suction side 20 is molded.

[0034] Shims (not shown) or other known tooling components for stiffening operations may then be applied to the surfaces of the blade 10 in preparation for stiffening (step 322). The shims may provide support and sealing for the composite component and enable the formation of a densified composite with a desired geometry.

[0035] Next, the unstiffened blade preform may be stiffened (step 326). The components may be stiffened at elevated temperatures and pressures. Although not so limited, the components may be stiffened at temperatures of about 200 to about 400°C and at pressures of about 3.45 bar (50 psig) to about 20.7 bar (300 psig). Additionally or alternatively, stiffening may include curing (e.g., heating), compression molding, bladder forming, or other suitable methods for hardening the blade assembly 100. Stiffening may include applying BN and SiC coatings using a chemical vapor infiltration (CVI) process, as is known in the art, thereby forming a rigid coated turbine blade preform.The term "co-stiffening" (and grammatical equivalents thereof) involves stiffening at substantially the same time, or at least during an overlapping period, during which two objects are stiffened. Co-stiffening can produce a substantially continuous matrix phase with additional strength, which is believed (but not intended to be limited by any theory) to be provided by an enhanced bond between the airfoil, dovetail, and integral platform. The term "unstiffened" (and grammatical equivalents thereof) describes objects that have not been stiffened at all, or at least objects that have been partially stiffened to a point such that the stiffening is immaterial.The term "partially stiffen" (and grammatical equivalents thereof) includes stiffening to a recognizable point, but not stiffening to a fully stiffened point. The term "fully stiffened" encompasses stiffening to a point for which an object is stiffened to a desired endpoint. The stiffening terms form a hierarchy with some overlap between adjacent terms. For example, the terms "unstiffened," "partially stiffened," and "fully stiffened" (with some overlap) express increasing amounts of stiffening. In an exemplary embodiment, the airfoil and dovetail preform and the integral platform preform may be co-stiffened with an initial partial stiffening followed by a subsequent stiffening.In all embodiments, a stiffened blade preform is formed when the stiffening is substantially fully completed.

[0036] In one embodiment, additional plies and / or angel wing preform structures may be applied to the platform surface 24. In one embodiment, an additional stiffening step as described above may be included. In yet another embodiment, additional plies may then be applied, followed by yet another stiffening step as described above. In an alternative embodiment, the blade replacement element may then be removed. In another alternative embodiment, the blade assembly 100 may then be arranged over a unitary dovetail and airfoil component, and additional plies may then be applied, followed by another stiffening step.

[0037] Next, the stiffened blade preform ring may be partially densified, for example, by introducing a carbonaceous slurry, as is known in the art, into the porosity of the stiffened blade preform (step 328). Finally, the stiffened blade preform may be further densified, for example, with at least silicon and alternatively boron-doped silicon, through a melt infiltration process that forms the blade, as is known in the art (step 328). Other techniques for forming components according to the present disclosure include polymer infiltration and pyrolysis (PIP). In this process, silicon carbide fiber preforms are infiltrated with a preceramic polymer, such as a polysilazane, and then heat treated to form a SiC matrix. Alternatively, the components may comprise an oxide / oxide process.In this type of processing, aluminum or aluminosilicate fibers can be pre-impregnated and then laminated into a pre-selected geometry and subsequently heated to form the ceramic matrix. Components can also be made from a carbon fiber reinforced silicon carbide (C / SiC) matrix (CMC). C / SiC processing involves a carbon fiber preform stacked in the pre-selected geometry. When used in a slurry casting process for SiC / SiC, the tool is made of a graphite material. The fiber preform is held at a temperature of approximately 1200°C during a chemical vapor infiltration process through the tool, forming the C / SiC CMC component.

[0038] The blade may then be machined (step 332) to produce the desired final geometry. In this embodiment, the platform 16 may include SiC-coated fibers and a polymer-based matrix. Materials such as a low-melting alloy, machining wax, and / or polymer materials may be used to coat the platform 16. To avoid adsorption of contaminants present in certain machining fluids, the blade 10 may be water-cooled during machining. The cutting direction may be preselected to avoid tearing out fiber materials. The cutting feed and speeds may also be predetermined to avoid damage to the blade 10 in the form of delamination or surface fiber removal.

[0039] In another embodiment, e.g., in CMC forming processes involving SiC / SiC slurry casting, C / SiC, or PIP, the process includes a tool made of graphite or other material suitable for use at temperatures above 816°C (1500°F). As described in Fig. 9, the tool is prepared (step 304), and the airfoil, dovetail, and two platforms are placed on top (steps 304, 306, and 318). However, for this embodiment, the removal window (mold or element 212) is not required. The tool 200 has a pocket on the suction side for the suction side platform. As in the Fig. 9, a bridge 214 is used for the pressure side platform 16. Thereafter, the preform is stiffened (step 326). For C / SiC and slurry-cast SiC / SiC, chemical vapor infiltration (CVI) is the first stiffening process. The tool 200 would be removed after part of the CVI process. For polymer infiltration and pyrolysis (PIP) component formation, the preform is infiltrated with a preceramic polymer, such as a polysilazane. The stiffening step in this embodiment is a heat treatment to convert the polymer to SiC or Si3N4. For the PIP process, the tool would be removed after several cycles. As in the Fig. 9, the compaction (step 328) and the processing (332) are then completed.

[0040] While the disclosure has been described with reference to a preferred embodiment, it will be recognized by those skilled in the art that various modifications could be made and equivalents could be substituted for the elements of the disclosure without departing from the scope of the disclosure. Furthermore, numerous changes could be made to adapt a particular situation or material to the teachings of the disclosure without departing from the essential scope thereof. Therefore, it is intended that the disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this disclosure, but that the disclosure will include all embodiments falling within the scope of the appended claims.

Claims

[1] A method of manufacturing a ceramic matrix composite blade (10), comprising: Placing at least one layer of a plurality of fiber ceramic layers in a preselected arrangement on a tool support surface (206) to form a blade and dovetail preform; Laying at least one further layer of the plurality of fiber ceramic layers onto the airfoil and dovetail preform in a second preselected arrangement to form an integral platform preform, wherein a plurality of the further fiber ceramic layers are stacked to form an unstiffened blade preform together with the blade and dovetail preform; stiffening the unstiffened blade preform to form a stiffened blade preform, wherein the airfoil (12), the dovetail (14) and the integral platform (16) comprise a substantially continuous co-stiffened matrix phase; and Compacting the stiffened blade preform to form the ceramic matrix composite blade (10) having an integral platform structure. [2] The method of claim 1, wherein the stiffening comprises partial stiffening and final stiffening. [3] The method of claim 1, wherein the compacting includes partially compacting the stiffened blade preform with a carbonaceous slurry. [4] The method of claim 1, wherein the densifying further includes densifying the stiffened blade preform with at least silicon. [5] The method of claim 1, wherein the stiffening and densifying are a slurry casting melt infiltration process. [6] The method of claim 1, wherein the stiffening and densifying are a chemical vapor phase infiltration process. [7] A method according to claim 1, wherein the stiffening and densification are a polymer infiltration and pyrolysis process. [8] The method of claim 1, further comprising applying a matrix material to the unstiffened blade preform. [9] The method of claim 1, further comprising applying at least one further fiber ceramic layer to form an angel wing preform after said partial stiffening and before said final stiffening. [10] The method of claim 8, further comprising adding one or more intermediate layers to the unstiffened blade preform. [11] The method of claim 8, further including machining the ceramic matrix composite blade (10) after compaction. [12] The method of claim 1, wherein the ceramic matrix composite blade (10) comprises pins. [13] Tool (200) for producing a ceramic matrix composite blade (10), comprising: a first tool component (202, 204, 212) having a first support surface (206) adapted to receive at least one ceramic fiber layer to form an airfoil and dovetail preform; a second tool component (214) having a second support surface (208) adapted to receive at least one further ceramic fiber layer to form an integral platform preform; wherein the first support surface (206) is selectively configurable to enable the integral platform preform to be laid against the airfoil and dovetail preform to form a blade preform, the first support surface having a dovetail shape (212) that is removable to alter the first support surface. [14] The tool (200) of claim 13, wherein the presence or absence of the dovetail shape (212) in the first tool component (202, 204, 212) determines the geometry of the first bearing surface (206). [15] Ceramic matrix composite blade assembly comprising: an airfoil (12), a dovetail (14) and an integral platform (16) having fiber reinforcement in a preselected arrangement in a ceramic matrix, the integral platform (16) including platform surfaces (24) and dovetail surfaces (26) enclosing the integral platform (16); wherein the airfoil (12), the dovetail (14) and the integral platform (16) all comprise a substantially continuous co-stiffened matrix phase. [16] The ceramic matrix composite blade assembly of claim 15, further comprising pins (30) disposed on the ceramic matrix composite blade assembly. [17] The ceramic matrix composite blade assembly of claim 15, further comprising angel wings (32) on the ceramic matrix composite blade assembly. [18] A ceramic matrix composite blade assembly according to claim 15, further comprising a separate member arranged and configured to seal a first blade from a second blade.

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