Patent Publication Number: US-2021163365-A1

Title: Needled ceramic matrix composite cooling passages

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application is a divisional of U.S. patent application Ser. No. 15/863,160, filed Jan. 5, 2018. 
    
    
     BACKGROUND 
     The present disclosure relates to ceramic matrix composite components, and more particularly, to an internal cooling passage. 
     Gas turbine engines typically include a compressor section to pressurize airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. Gas path components, such as turbine blades, often include airfoil cooling that may be accomplished by external film cooling, internal air impingement, and forced convection, either separately, or in combination. 
     Ceramic matrix composite (CMC) components can endure high temperatures, but those temperatures may be below the gas path operational temperatures of some modern turbine engine stages. Internal convective cooling of CMC components may be primarily from impingement baffle like structures and film cooling from cooling passages. The cooling passages are either drilled using laser or Electron Discharge Machining (EDM). Laser passages are ideal for relatively short passages such as those at the leading edge of an airfoil, while EDM is ideal for long passages such as those at the trailing edge. EDM relies on a current supplied to an electrode which is discharged through a grounded part, however CMC material cannot carry current, and cannot use EDM. Machining cooling passages in the CMC component may result in cut fibers in the CMC material which may weaken the CMC component or expose a surface to environmental attacks. 
     SUMMARY 
     A method for forming a hole within a ceramic matrix composite component according to one disclosed non-limiting embodiment of the present disclosure includes forming a first core portion for a ceramic matrix composite component; embedding a hollow member into the first core portion at a desired location; wrapping the first core portion with a first ceramic matrix composite material; inserting a rod through the hollow member and into the first core portion; removing the hollow member; assembling a second core portion to the first core portion such that the rod extends into the second core portion; and wrapping the first core portion and the second core portion with a second ceramic matrix composite material. 
     A further aspect of the present disclosure includes separating a plurality of fibers around the hollow member. 
     A further aspect of the present disclosure includes penetrating the ceramic matrix composite material with the hollow member. 
     A further aspect of the present disclosure includes coaxially forming a hole in the first core portion for the rod within a hole in the first core portion for the hollow member. 
     A further aspect of the present disclosure includes removing the hollow member and leaving the rod in the first core portion. 
     A method for forming a hole within a ceramic matrix composite component according to one disclosed non-limiting embodiment of the present disclosure includes forming a first core portion for a ceramic matrix composite component; embedding a hollow member into the first core portion at a desired location; wrapping the first core portion with a first ceramic matrix composite material; inserting a rod through the hollow member and into the first core portion; assembling a second core portion to the first core portion such that the rod extends into the second core portion; and wrapping the first core portion and the second core portion with a second ceramic matrix composite material; and burning out the first core portion, the second core portion, the hollow member, and the rod from the ceramic matrix composite material subsequent to wrapping the first core portion and the second core portion with the second ceramic matrix composite material. 
     A further aspect of the present disclosure includes that the hollow member is manufactured of a nylon. 
     A core for a ceramic matrix composite component according to one disclosed non-limiting embodiment of the present disclosure includes a first core portion; a second core portion; and a rod extending between the first core portion and the second core portion. 
     A further aspect of the present disclosure includes that the first core portion is adjacent a trailing edge of the ceramic matrix composite component. 
     A further aspect of the present disclosure includes that the second core portion is adjacent a leading edge of the ceramic matrix composite component. 
     A further aspect of the present disclosure includes that the second core portion will form an impingement passage within the ceramic matrix composite component. 
     A further aspect of the present disclosure includes that the rod is manufactured of the same material as the first core portion and the second core portion. 
     A further aspect of the present disclosure includes that the rod is of a desired cooling hole shape. 
     A further aspect of the present disclosure includes that the rod is glued into the first core portion. 
     A further aspect of the present disclosure includes that the first ceramic matrix composite material forms an internal wall within an airfoil. 
     A further aspect of the present disclosure includes a plurality of fibers through which the rod extends but does not cut. 
     A further aspect of the present disclosure includes a second ceramic matrix composite material that wraps the first core portion and the second core portion. 
     A further aspect of the present disclosure includes a hole for the rod coaxial within a hole for a hollow member. 
     A further aspect of the present disclosure includes that an inner surface of the hollow member is sized to receive the rod, the hollow member extending between the first core portion and the second core portion. 
     A further aspect of the present disclosure includes that each of the first core portion and the second core portion include a multiple of grooves. 
     The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be appreciated; however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of an example gas turbine engine architecture; 
         FIG. 2  is an exploded view of rotor assembly with a single representative ceramic matrix composite turbine blade; 
         FIG. 3  is a cross-sectional illustration of an example ceramic matrix composite turbine blade of the gas turbine engine. 
         FIG. 4  is a cross-sectional illustration of an example ceramic matrix composite turbine blade taken along line  4 - 4  in  FIG. 3 . 
         FIG. 5  is a cross-sectional illustration of an example ceramic matrix composite turbine blade taken along line  4 - 4  in  FIG. 3  illustrating a method to form a passage in the ceramic matrix composite component. 
         FIG. 6  illustrates a flow diagram of an example method to form a passage in the ceramic matrix composite component. 
         FIG. 7  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 6  illustrating drilling a counter bored hole. 
         FIG. 8  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 6  illustrating embedding a hollow member into the core. 
         FIG. 9  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 6  illustrating wrapping with a ceramic material such that the hollow member penetrates therethrough. 
         FIG. 10  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 6  illustrating inserting a rod into the hollow member. 
         FIG. 11  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 6  illustrating removing the hollow member. 
         FIG. 12  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 6  illustrating the ceramic material closing around the rod. 
         FIG. 13  illustrates a flow diagram of an example method to form an internal passage in the ceramic matrix composite component. 
         FIG. 14  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating drilling a counterbored hole in one core portion. 
         FIG. 15  illustrates a perspective view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating embedding a hollow member into the core portion. 
         FIG. 16  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating wrapping the core portion with a ceramic material such that the hollow member penetrates therethrough. 
         FIG. 17  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating inserting a rod into each hollow member. 
         FIG. 18  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating removing the hollow member from an interface surface of the core portion. 
         FIG. 19  illustrates a perspective view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating assembling two core portions. 
         FIG. 20  illustrates a cross-sectional view of a core for a ceramic matrix composite component illustrating a step of the method of  FIG. 13  illustrating assembling two example core portions. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  as disclosed herein is a two spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26 , and a turbine section  28 . The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26 , then expansion through the turbine section  28 . Although depicted as a high bypass gas turbofan engine architecture in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein are not limited only thereto. 
     The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation around an engine central longitudinal axis A relative to an engine case structure  36  via several bearings  38 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor (“LPC”)  44  and a low pressure turbine (“LPT”)  46 . The inner shaft  40  drives the fan  42  directly or through a geared architecture  48  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor (“HPC”)  52  and high pressure turbine (“HPT”)  54 . A combustor  56  is arranged between the HPC  52  and the HPT  54 . 
     With reference to  FIG. 2 , a rotor assembly  60  such as a turbine rotor assembly includes an array of blades  84  (one shown) circumferentially disposed around a disk  86 . Each blade  84  includes a root  88 , a platform  90  and an airfoil  92 . The blade root  88  is received within a rim  94  of the disk  86  and the airfoil  92  extends therefrom. 
     The platform  90  separates a gas path side inclusive of the airfoil  92  and a non-gas path side inclusive of the root  88 . The airfoil  92  defines a blade chord between a leading edge  98 , which may include various forward and/or aft sweep configurations, and a trailing edge  100 . A first sidewall  102  that may be convex to define a suction side, and a second sidewall  104  that may be concave to define a pressure side are joined at the leading edge  98  and at the axially spaced trailing edge  100 . The tip  96  extends between the sidewalls  102 ,  104  opposite the platform  90 . 
     With reference to  FIG. 3 , to resist the high temperature stress environment in the gas path of a turbine engine, each blade  84  includes an array of internal passageways  110 . The array of internal passageways  110  includes a multiple of feed passages  112  through the root  88  that communicates airflow into a multiple of cavities  114  (shown schematically) within the airfoil  92 . The cavities  114  distribute the cooling flow through passages  130  in the sidewalls  102 ,  104 , leading edge  98 , and/or the trailing edge  100  (also shown in  FIG. 4 ). Impingement passages  132  (FIG. 4 ) may also be located though internal walls  134  between one or more of the array of internal passageways  110 . It should be appreciated that various feed architectures, cavities, and passageway arrangements will benefit herefrom. 
     With reference to  FIG. 5 , the example cooled turbine airfoil  84  is manufactured as a CMC component. Though the CMC may have less strength relative to metallic counterparts, CMCs can endure higher material temperatures and are significantly lighter. Although a turbine blade will be used to illustrate the disclosed cooling passage formation method, other components will also benefit herefrom. 
     The example turbine airfoil  84  is generally formed from a core  200  which may be formed from multiple portions  200 A,  200 B which are wrapped with a ceramic material  202 . The core  200  is later removed such that the cured ceramic material  202  forms the airfoil  92  and the array of internal passageways  110 . The core  200  may comprise a material such as carbon. The core  200  is readily cast and/or machined with conventional methods then later removed without damage to the ceramic material  202 . The core  200  may include a multiple of longitudinal grooves  201 . 
     The ceramic material  202  may be an arrangement of ceramic fibers  204 . Examples of the ceramic material  202  may include a three-dimensional weave of the ceramic fibers  204 . Alternatively, or in addition, the ceramic material  202  may include a two-dimensional weave of the ceramic fibers  204 . The ceramic material  202  may include multiple layers of two-dimensional weave of the ceramic fibers  204 . Alternatively, or in addition, the ceramic material  202  may include a fiber layup, such as a unidirectional layup. In some examples, each of the ceramic fibers  204  may be a bundle and/or a tow of ceramic fibers. The fibers in each bundle or tow may be braided or otherwise arranged. The ceramic fibers  204  may comprise a material that is stable at temperatures above  1000  degrees Celsius. Examples of the ceramic fibers  204  may include fibers of alumina, mullite, silicon carbide, silicon, zirconia or carbon. 
     With reference to  FIG. 6 , a method  300  for forming the passages  130  through, for example, the airfoil sidewall  102 ,  104  ( FIG. 4 ) in a ceramic matrix composite component is illustrated in a schematic block diagram form. It should also be appreciated that application is not limited to aerospace components and various other applications will benefit herefrom. 
     Once the core  200  is manufactured ( 302 ), a counterbored hole  212 / 216  ( FIG. 7 ) is drilled ( 304 ;  FIG. 6 ) at each location in which the passages  130  are to be formed. The counterbored hole  212 / 216  includes a blind hole  212  for a rod  214  within a blind hole  216  for a hollow member  218  along a common axis  220  ( FIG. 5 ). A step  222  is formed between the blind holes  212 ,  216  to form a stop for the hollow member  218 . Each counterbored hole  212 / 216  is located and oriented to form the respective passages  130 . 
     Next, the hollow member  218  is located in each blind hole  216  ( 306 ,  FIG. 6 ;  FIG. 8 ). The hollow member  218  has an interior diameter equal to or greater to the size of the desired passage  130 . The outside diameter may be equal to or slightly smaller than the blind hole  216 . The hollow member  218  may include a sharp end  219  to form a hollow needle and may be manufactured of a metal alloy, a nylon, or any other rigid material that is compatible with the CMC material. 
     Next, the core  200  is wrapped with the ceramic material  202  using the hollow member  218  to pierce through the ceramic material  202  ( 308 ,  FIG. 6 ;  FIG. 9 ). The ceramic material  202  comprises the plurality of fibers  204  through which the hollow member  218  extends but does not cut. The hollow members  218  are of a strength to penetrate and separate the plurality of fibers  204 . 
     The core  200  is wrapped with the ceramic material  202  to form a ceramic matrix composite body that may be the CMC component in which the passage  130  is to be formed. Alternatively, the ceramic matrix composite body may be a component of the CMC component in which the passage  130  is to be formed. The ceramic matrix composite body may comprise of, for example, a silicon carbide ceramic matrix composite. The ceramic matrix composite body may have any shape or form, not just the shape illustrated. Once all the layers of the ceramic material  202  are in place, one rod  214  is inserted into each hollow member  218  ( 310 ,  FIG. 6 ;  FIG. 10 ). 
     The rod  214  is shaped and sized to form the desired passages  130 . The rod  214  may be formed of the same material as the core such as a carbon. The rod  214  may be circular, rectilinear, oval, racetrack, or of other cross-sectional shape. Optionally, each rod  214  may be glued into each blind hole  212  with a glue  223 . 
     Next, the hollow member  218  is removed ( 312 ,  FIG. 6 ;  FIG. 11 ) leaving the rod  214  in place. The ceramic material  202  then closes ( FIG. 12 ) around the rod  214 . The ceramic fibers  204  of the ceramic material  202  are not broken in this process, such that it is readily apparent that this method was used because any drilling method would result in the cutting of the ceramic fibers  204 . Alternatively, the hollow member  218  can be manufactured of a material such as a nylon which can be readily burned out with the core  200  and then the rod  214  is burned out so that the hollow member need not be removed. That is, the hollow member  218  burns out at a lower temperature than infiltration temps then the rod  214  burns out with the core  200 . 
     Next, the ceramic material  202  is cured ( 314 ,  FIG. 6 ) per conventional CMC manufacturing procedures to form the CMC component. Forming the cooled turbine airfoil  84  as the CMC component may include infiltrating a molten metal or alloy into the ceramic material  202 . The multiple of longitudinal grooves  201  ( FIG. 7 ) facilitate the infiltration. 
     The molten metal or alloy fills the gaps between the ceramic fibers  204  and the rods  214 . The molten metal or alloy may also react with a reactive element source present in the ceramic material  202  to form additional silicon based ceramic matrix material. In some examples, a chemical vapor infiltration coating may be applied to the ceramic material  202  prior to the melt infiltration to stiffen the ceramic fibers  204 . Alternatively, or in addition, forming the CMC component from the ceramic material  202  may include chemical vapor infiltrating the ceramic material  202  instead of melt infiltrating. 
     Finally, the core  200  and rods  214  are removed ( 316 ,  FIG. 6 ) via heat, acid, or other method which does not harm the ceramic material  202  per conventional CMC manufacturing procedures. Once the core  200  and rods  214  are removed, the passages  130  and the array of internal passageways  114  are formed. 
     With reference to  FIG. 13 , a method  400  for forming the internal passages  132  through, for example, the internal walls  134  ( FIG. 4 ) in a ceramic matrix composite component is illustrated in a schematic block diagram form. It should also be appreciated that application is not limited to aerospace components and various other applications will benefit herefrom. 
     Once the core portions  200 A,  200 B are manufactured ( 402 ,  FIG. 13 ), the counter bored holes  210  are drilled ( 404 ,  FIG. 13 ;  FIG. 14 ) at each location in which the external passages  130  and impingement passages  132  are to be formed. Although the first core portion  200 A and the second core portion  200 B are illustrated to represent formation of the core  200 , any number of core portions may be utilized to form a desired internal structure. 
     Next, one hollow member  218  is located in each of the blind holes  216  ( 406 ,  FIG. 13 ;  FIG. 15 ) including those that are in an interface surface  230 B. The interface surfaces  230 A,  230 B are shaped to form one side of the internal wall  134  that separates the cavities  114 . 
     Next, at least one core portion, here represented by core portion  200 B, is wrapped ( 408 ,  FIG. 13 ;  FIG. 16 ) with the ceramic material  202  using the hollow member  218  to pierce through the ceramic material  202  as discussed above. 
     Next, the rods  214  are inserted in each of the respective hollow members  218  ( 410 ,  FIG. 13 ;  FIG. 17 ). 
     Next, the hollow members  218  are removed ( 412 ,  FIG. 13 ;  FIG. 18 ) from the interface surface  230 B leaving only the rods  214  in place. That is, the hollow members  218  remain in all surfaces other than the interface surface  230 B which will interface with the interface surface  230 A of core portion  200 A. The core portions  200 A,  200 B may still retain the embedded hollow members  218  in surfaces other than the interface surfaces  230 A,  230 B such as those that will form surfaces of the component such as the airfoil sidewalls  102 ,  104 . 
     Next, the core portion  200 A is assembled ( 414 ,  FIG. 13 ;  FIG. 19 ) to the core portion  200 B. The core portion  200 A includes a multiple of blind holes  212  that correspond with each of the rods  214  in the core portion  200 B. The rods  214  extend from the interface surface  230 B to span the core portion  200 A and the core portion  200 B with the ceramic material  202  trapped therebetween. The rods  214  may be glued in place to retain the core portions  200 A, such that the ceramic material  202  that forms the internal walls  134  is compressed between the core portions  200 A,  200 B to form a desired shape. That is, the interface surfaces  230 A,  230 B may be shaped to form the internal walls  134  into a desired shape. 
     Next, a multiple of core portions, here shown as the core portions  200 A,  200 B, are wrapped ( 416 ,  FIG. 13 ;  FIG. 20 ) with the ceramic material  202  using the hollow members  218  to pierce through the ceramic material  202  as described above. Any number of core portions may be respectively assembled to form the core  200  with particular core portions being wrapped as in  408 . 
     The method  400  then continues with forming of the ceramic material  202  to form the ceramic matric composite component, then removal of the core portions and rods as described above in accordance with method  300 . 
     The “cast in” passages  130 ,  132  are readily identifiable, may be of various cross-sectional shapes, reduce machining time, and facilitate the manufacture of long passages through CMC components such as those through the trailing edge of an airfoil. 
     The use of the terms “a”, “an”, “the”, and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier “about” used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. 
     Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.