PRODUCTION OF CAVITIES IN COMPONENTS MADE FROM POROUS SUBSTRATES

Description

BACKGROUND OF THE INVENTION

[0001] This application relates to a way of forming voids, such as cavities or hollows, or cooling channels, into a substrate material in ceramic matrix composite ("CMC").

[0002] Components are being manufactured from more sophisticated materials. One such material includes CMC materials. Such materials have very beneficial characteristics when facing certain operational situations.

[0003] As one example, CMC components are very resistant to heat and, thus, have many applications in high temperature environments. They are being utilized in any number of high temperature locations in gas turbine engines, as an example. However, some applications would benefit from the provision of voids within the CMC body for various reasons. One reason may be the provision of cooling air through channels.

[0004] However, the formation of hollows, channels or other voids within a CMC body has been challenging.

[0005] It has been proposed to dispose carbon strings within a substrate and then oxidize the carbon away, leaving a void. However, these are very large diameter voids.

[0006] A method having the features of the preamble of claim 1 is disclosed in US 2005 / 118392 A1.

[0007] US 2014 / 072736 A1 discloses a method of producing an internal cavity in a ceramic matrix composite and mandrel therefor.

[0008] US 2013 / 017094 A1 discloses a method of fabricating a composite material blade having internal channels, and a composite material turbine engine blade.

[0009] US 2002 / 084558 A1 discloses a SiC-C / C composite material, uses thereof, and method for producing the same.

[0010] US 5196271 A discloses a method of making ceramic articles having channels therein and articles made thereby.SUMMARY OF THE INVENTION

[0011] According to the present invention, there is provided a method of creating a component as claimed in claim 1.

[0012] The substrate material may initially include graphite, such that the matrix includes a silicon carbide material.

[0013] The voids may be utilized to receive a material subsequent to the formation of the voids.

[0014] The voids may be subsequently filled.

[0015] The template material may be one of fibers, particles, or sheets.

[0016] These and other features may be best understood from the following drawings and specification.BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 schematically shows a gas turbine engine. Figure 2 shows a turbine blade which may incorporate the teachings of this disclosure. Figure 3A shows an initial step according to the method of this disclosure. Figure 3B shows a subsequent step. Figure 4A shows an initial step in another embodiment. Figure 4B shows the subsequent step. Figure 5A shows yet another embodiment. Figure 5B shows a subsequent step. Figure 6 shows another method embodiment. DETAILED DESCRIPTION

[0018] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

[0019] The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.

[0020] The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

[0021] The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.

[0022] The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

[0023] A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition -- typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. "Low corrected fan tip speed" is the actual fan tip speed in ft / sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)] 0.5< . The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second (350.5 meters / second).

[0024] Figure 2 shows a component body 120, which may be a turbine blade as utilized in a gas turbine engine such as the engine in Figure 1. Of course, other components such as heat panels, vanes, combustor liners, seals and the like may benefit from the teachings of this application.

[0025] As known, a source of air 122 provides cooling air into an airfoil 124 formed on the component 120. As shown schematically in Figure 2, there are a plurality of very small voids, cavities or channels 126 extending through the component 120 and which may be formed by the teachings of this disclosure.

[0026] Figure 3A shows a first intermediate step to form the component 120. A carbon-containing substrate 130 is provided with silicon template materials 132. In non-limiting examples, the carbon-containing substrate can be graphite, amorphous carbon, glassy carbon, carbonaceous material, activated carbon, turbostratic carbon and mixtures thereof. The silicon particle may be of any form including fibers, whiskers, sheets, etc. and in very small sizes.

[0027] The matrix is then processed via typical processing. As an example, polymer infiltration and pyrolysis, slurry or tape casting, etc. may be performed to further process the matrix. This processing occurs at a temperature below the melting point of the silicon template material 132.

[0028] After a prescribed state has been reached, as an example, an appropriate porosity level for substrate 130, the matrix is put through a melt infiltration step at a temperature of above the melting point of the template material 132. Thus, the template will melt and wick or infiltrate or otherwise relocate into the porosity of the substrate and leave a void space such as shown at 126 in Figure 3B. Thus, the voids 126 become the cooling channels, in this example, while the matrix 134 provides the body of the component 120. Appropriate choice of substrate and template morphologies enables control of the geometries of voids 126. As an example, fibrous substrates and particulate templates can produce voids 126 with an elongated morphology.

[0029] This method is applicable to components formed by other materials, including other composites or monolithic structures or organic matrix composites. Notably, particular attention may be required with regard to the template materials for these alternative matrix materials.

[0030] This method, thus, forms cooling channels, hollow cavities, or other deliberately formed voids, which can be costly and difficult to produce by other methods within CMC materials. Challenges to incorporating these features from the beginning of CMC processing include design complexity due to the ceramic fiber architecture and compatibility issues with subsequent CMC matrix processing methods. As an example, many processes would fill in the desired voids or hollows.

[0031] Challenges with forming hollow features subsequent to CMC processing primarily include high cost and technical limitations of machining and may result in damage to the CMC matrix or fiber. As an example, loss of strength or environmental durability due to fabric or matrix damage can occur. Also, additional processing treatments can raise challenges with regard to maintaining the voids in the matrix.

[0032] This in situ process to form the voids overcomes traditional design barriers. In particular, the cost of this method is relatively low and the likelihood of damage is also reduced from the prior art.

[0033] Other benefits with regard to CMC include facile incorporation of the invention with typical CMC materials.

[0034] Also, there may be remnant melt infiltration materials that are targeted. In the example component 123 shown in Figure 4A, the substrate 140 is a carbon-rich polymer derived ceramic. The template is again silicon 142.

[0035] As shown in Figure 4B, after the melt process, the matrix 144 remains with the silicon template now having infiltrated the porosity of the matrix 144, leaving voids 146. As shown, there may be a layer 149 of silicon around the void or opening 146. This can have additional environmental benefits as it isolates, or insulates, at least in part, the cooling air passing through the passage 146 from contacting the CMC material.

[0036] Figure 5A shows another embodiment 160 wherein there is a silicon template 162 and a substrate 164. After undergoing melting (see Figure 5B), voids 168 remain within the matrix 166. The substrate 164 may again be a carbon-rich polymer derived ceramic material, however, in another example, relatively small molybdenum, or other metallic particles 161 may also be utilized. This can result in molybdenum disilicide, or other metal silicide, regions 170 in the final matrix 166 as shown in Figure 5B, preferentially reducing or eliminating the original silicon template material. The silicon template material can be partially or wholly eliminated in the final form. The MoSi 2 , or metal silicide, which is formed has a much higher melting point material than the original Si. While molybdenum is disclosed, metallic particles of other metals, including titanium, vanadium, chromium, cobalt, nickel, zirconium, niobium, hafnium, tantalum, tungsten, platinum, yttrium and boron may be utilized. All of these materials would result in a metal silicide region as mentioned above with regard to the molybdenum embodiment.

[0037] Figure 6 schematically shows a subsequent method step wherein a tool 204 is depositing infiltration material 206 into or onto a component 202 having voids 200 formed by the method of this disclosure. Thus, this method may be utilized to form such voids as an intermediate step such that the voids 200 better facilitate the depositing or infiltration of material 206 throughout the component 202. The voids may then be utilized for subsequent deposition or infiltration processes, to allow material to pass through the voids, and then into the matrix. In some of these methods, the voids can remain hollow after the subsequent process. Additionally, this method can be utilized not only to have voids in a final component, but may be utilized to form such voids as an intermediate step, the voids 202 then being filled by different materials for various known reasons.

[0038] In this invention, a diameter of the void is extremely small. Average hydraulic diameters are greater than or equal to about ten and less than or equal to about 20 microns in scale.

[0039] The teachings are broadly applicable to hybrid CMC processing methods including polymer infiltration and pyrolysis, chemical vapor infiltration, melt infiltration, and glass transfer molding as examples.

[0040] Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the following claims

Claims

1. A method of creating a component (120, 202) comprising: forming a ceramic matrix composite (CMC) substrate (130) and depositing a template material (132) within said substrate (130), such that there are a plurality of template members; processing the substrate (130) to provide an appropriate porosity level; and heating said substrate (130) to a temperature above a melting point of said template material (132), such that said template material (132) wicks into the porosity of said substrate (130) and forms a component (120, 202) formed of a matrix material (134) having voids (126), wherein an average hydraulic diameter of said voids (126, 200) is greater than or equal to about 10 and less than or equal to about 20 microns, wherein said template material (132) is silicon and said component (120, 202) is for use in a gas turbine engine.

2. The method as set forth in claim 1, wherein said substrate (130) initially includes graphite, such that said matrix material includes a silicon carbide material.

3. The method as set forth in claim 1 or 2, wherein said voids (126) are utilized to receive a material subsequent to the formation of the voids.

4. The method as set forth in any preceding claim, wherein said voids (126, 200) are subsequently filled.

5. The method as set forth in any preceding claim, wherein said template material is one of fibers, particles, or sheets.