Patent Publication Number: US-10787914-B2

Title: CMC airfoil with monolithic ceramic core

Description:
BACKGROUND 
     The disclosed subject matter relates generally to nonmetallic airfoils and more particularly to ceramic airfoils. 
     Laminated ceramic matrix composite (CMC) airfoils are well known for gas turbine engines, but have certain shortcomings Though extremely light in weight and exhibiting tolerance of foreign object damage (FOD), they are expensive to process into complex aerodynamic shapes. Conversely, ceramic airfoils are easier to form than laminated CMC airfoils, but are prone to large scale fracture due to FOD. 
     Attempts have been made to produce a reliable hybrid ceramic/CMC airfoil. However, it is difficult to combine a CMC shell with a ceramic spar due to limited ways of joining the two materials. Further, when using traditional CMC processing steps, large portions of the CMC have to contact the ceramic spar in order to accurately form the airfoil surfaces. This leaves little or no room for spaces or passages between the spar and shell, for example, to provide cooling air to the spar without sacrificing the smoothness of the CMC airfoil surface. It also requires the ceramic of the spar and the ceramic matrix of the shell to have closely matched chemical, mechanical, and thermal properties at elevated temperatures to avoid damaging chemical reactions and/or residual stress. 
     SUMMARY 
     An airfoil comprises a core having a first surface, a skin having a second surface disposed over at least a portion of the first surface of the core, and at least one of a transient liquid phase (TLP) bond and a partial transient liquid phase (PTLP) bond. The at least one bond is disposed between the first surface and the second surface, joining the skin to the core. 
     A method for making a hybrid airfoil component comprises providing a ceramic airfoil core. A ceramic matrix composite (CMC) airfoil skin is placed over at least a portion of the ceramic airfoil core. The CMC skin is joined to the ceramic core to define an airfoil shape. The joining step is performed at least in part by forming a partial transient liquid phase (PTLP) bond between the ceramic core and the CMC skin. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  shows a gas turbine engine. 
         FIG. 2  is a portion of a rotor disk and a hybrid ceramic/CMC airfoil. 
         FIG. 3A  is a first sectional view taken across line  3 A- 3 A of the airfoil shown in  FIG. 2 . 
         FIG. 3B  is a second sectional view of the airfoil taken across line  3 B- 3 B of  FIG. 3A . 
         FIG. 4A  shows a first PTLP bond joining the suction side CMC skin to the adjacent ceramic core. 
         FIG. 4B  shows an example configuration setting up the first PTLP bond shown in  FIG. 4A . 
         FIG. 5A  depicts a first alternate configuration of an airfoil with PTLP bonds on either side of a thermal protection structure, which together join the CMC skin and the ceramic core. 
         FIG. 5B  is a second alternate configuration of an airfoil with a PTLP bond between two thermal protection elements forming a thermal protection structure joining the CMC skin and the ceramic core. 
         FIG. 6  shows steps of a method for making a hybrid ceramic/CMC airfoil. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  is a schematic view of gas turbine engine  20 . Gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates fan section  22 , compressor section  24 , combustor section  26  and turbine section  28 , although alternative gas turbine designs (including designs utilizing a power turbine in place of fan section  22 ) may also benefit from the described subject matter. In turbofan embodiments, fan section  22  drives air along bypass flowpath B, while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26 , and then expansion through the turbine section  28 . 
     Dual-spool embodiments such as example engine  20  generally include low-speed spool  30  and high-speed spool  32  mounted for rotation about an engine central longitudinal axis A. Spools  30 ,  32  rotate relative to engine static structure  36  via several bearing systems  38 . It should be understood that different numbers of spools, as well as various bearing systems  38  may alternatively or additionally be provided. 
     Low-speed spool  30  generally includes inner shaft  40  that interconnects a fan  42 , low-pressure compressor  44  and low-pressure turbine  46 . In certain turbofan embodiments, inner shaft  40  can be connected to fan  42  through geared architecture  48  to drive fan  42  at a lower speed than low-speed spool  30 . High-speed spool  32  includes outer shaft  50  that interconnects high-pressure compressor  52  and high-pressure turbine  54 . Combustor  56  is arranged between high-pressure compressor  52  and high-pressure turbine  54 . Mid-turbine frame  57  of the engine static structure  36  can be arranged axially between high-pressure turbine  54  and low-pressure turbine  46 . Mid-turbine frame  57  can further support bearing systems  38  in turbine section  28 . Inner shaft  40  and 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. 
     The core airflow is compressed by low-pressure compressor  44  and then by high-pressure compressor  52 , mixed and burned with fuel in combustor  56 , then expanded over high-pressure turbine  54  and low-pressure turbine  46 . Combustor  56  is therefore in fluid communication with the compressor section, to receive air compressed by low-pressure compressor  44  and high-pressure compressor  52 . Mid-turbine frame  57  can also include airfoils  59  which are in the core airflow path. Turbines  46  and  54  are in fluid communication with combustor  56 , wherein the expanding gas provided by combustor  56  drives the respective low-speed spool  30  and high-speed spool  32 . 
       FIG. 2  shows a portion of gas turbine rotor assembly  62 , which includes rotor disk  64  with a plurality of circumferentially distributed hybrid rotor blades  66  (one shown in  FIG. 2 ). Hybrid rotor blade  66  includes airfoil section  68 , root section  70 , leading edge  72 , trailing edge  74 , pressure surface  76 , suction surface  78 , radial retention slots  80 , pressure-side root bearing surface  82 , disk bearing surfaces  84 , disk teeth  86 , forward bearing surface  88 , aft bearing surface  90 , retention ring  92 , and shim  94 . 
     Certain embodiments of rotor assembly  62  and/or hybrid rotor blade  66  are disposed in the hot section, such as high-pressure turbine  54 , or low-pressure turbine  46  as shown in  FIG. 1 . Additionally or alternatively, rotor assembly  62  may be disposed in fan section  22 , low-pressure compressor section  44 , and/or high-pressure compressor section  50 . In other alternative embodiments, hybrid airfoil sections can be formed in a similar manner for one or more stator assemblies in these sections of engine  20 . 
     In  FIG. 2 , airfoil section  68  can include leading edge  72 , trailing edge  74 , pressure surface  76 , and suction surface  78 . Root section  70  can be a single root with circumferentially opposed bearing surfaces for securing hybrid blade  66  into a corresponding radial retention slot  80  of disk  64 . Alternatively, root section  70  can be a multilobe root. In  FIG. 2 , pressure-side root bearing surface  82  and an opposing suction-side bearing surface (not visible) mate with respective bearing surfaces  84  of disk teeth  86 , which define a longitudinal extent of slot  80 . Root section  70  includes longitudinally facing forward bearing surface  88  and aft bearing surface  90  (not visible in  FIG. 2 ). At least one of these longitudinally facing bearing surfaces can be secured using one or more retention rings  92 , or alternatively using another bearing surface of the disk (not shown). Shim  94  can be disposed annularly between blade root section  70  and the corresponding radial retention slot  80 . 
     It will be recognized that certain embodiments of rotor assembly  62  can include an inner-diameter flow surface defined, for example, by a plurality of circumferentially distributed blade platforms. Such platforms may be integrally formed or secured to each hybrid blade  66  proximate the transition between airfoil section  68  and root section  70 . Likewise, certain embodiments may also include an outer-diameter flow surface that may be integrally formed or secured to each hybrid blade  66  proximate the outer tip of the airfoil. However, to better illustrate other elements, any possible inner or outer flow surface or blade platform has been omitted from the examples described herein. 
     As shown in more detail in  FIGS. 3A and 3B , hybrid blade  66  can include a hybrid airfoil section  68  in which a core with a first (e.g., ceramic) outer surface is bonded to a second (e.g., ceramic matrix composite/CMC) inner surface of an airfoil skin. The skin can be disposed over at least a portion of the outer surface of the ceramic core to define one or more airfoil surfaces such as pressure surface  76  and/or suction surface  78 . At least one of a transient liquid phase (TLP) bond and a partial transient liquid phase (PTLP) bond can be disposed between the first outer surface and second inner surface, thereby joining the CMC skin to the ceramic core to define a shape of airfoil section  68 . Due to reduced weight and moment of inertia, as well as the ability to form complex shapes, airfoil section  68  can be highly tapered to increase engine efficiency. 
       FIG. 3A  is a first sectional view taken across line  3 A- 3 A of the airfoil shown in  FIG. 2 .  FIG. 3B  is a sectional view taken across line  3 B- 3 B of  FIG. 3A , showing an example construction of hybrid blade  66  in more detail. 
     As seen in  FIG. 3A , airfoil section  68  of hybrid blade  66  generally includes ceramic core  96 , CMC skin portions  98 A,  98 B, and PTLP bonds  100 ,  102 . Suction-side CMC skin portion  98 A is joined to ceramic core  96  by one or more suction-side PTLP bonds  100 . Pressure-side CMC skin portion  98 B can be generally spaced from ceramic core  96  except proximate a location of one or more pressure-side PTLP bonds  102  and thermal protection structures  104 . This defines one or more thermal protection spaces  106  between thermal protection structures  104  and ceramic core  96  to reduce thermal conduction from hot gases impinging on pressure-side CMC skin portion  98 B. For example, the hot gases can be working gases when airfoil  68  is used in hot section and/or power turbine applications. Thermal protection spaces  106  can also serve as cooling passages and can be placed in communication with any cooling passages (not shown) which may be formed through ceramic core  96 . Thermal protection structures  104  and PTLP bond(s)  102  allow for greater differential thermal expansion between core  96  and CMC skin portion  98 B. Thus the respective ceramic materials in core  96  and CMC skin portions  98 A,  98 B can be selected with less concern of damage that can be caused by differential thermal growth. 
     The inner surface of the CMC skin can extend over some or all of the outer surface of the ceramic core. In the example shown, the CMC skin does not extend over the entirety of airfoil section  68 . As shown in  FIG. 3A , ceramic core  96  has leading-edge portion  108  defining airfoil leading edge  72 , as well as trailing-edge portion  110  defining airfoil trailing edge  74 . This configuration is shown in part because it allows for simple incorporation of CMC sheets to define substantial portions of pressure surface  76  and suction surface  78 . This configuration allows for CMC skin portions  98 A,  98 B to hold together ceramic core  96  in the event of failure (e.g., from a foreign object strike) while simplifying manufacture of the outer CMC surfaces and incorporation of the same into airfoil section  68 . However, it will be appreciated that a substantially contiguous CMC skin can also extend over some or all of leading edge  72  and trailing edge  74 , as well as the airfoil tip. 
     A hybrid blade also provides increased FOD resistance, especially in larger airfoils. Instead of potential perforation of a CMC blade, or failure of a ceramic blade, the energy absorption characteristics of ceramic core  96  and CMC skin portions  98 A,  98 B often will keep airfoil section  68  intact for a more graceful failure, which can prevent cascading foreign object damage to the engine. In any of these embodiments, the hybrid configuration also offers increased flexibility in the complexity of small details and complex shapes with monolithic ceramics relative to a CMC structure. Spaces  106  can also double as skin cooling passages depending on the configuration of thermal protection structures  104 . 
     Ceramic core  96  can be a monolithic ceramic, i.e., not reinforced by internal fibers or the like. However, core  96  can include cooling passages  111  formed during or after casting. In certain embodiments, ceramic core  96  includes at least one ceramic compound selected from one of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), silicon carbide (SiC), tungsten carbide (WC), and zirconium oxide (ZrO 2 ). 
     Suction- and pressure-side CMC skin portions  98 A,  98 B can be individually or integrally formed from a plurality of fibers disposed in a ceramic matrix. Example fibers can include combinations of silicon carbide (SiC), titanium carbide (TiC), aluminum oxide (Al 2 O 3 ), and/or carbon (C). The ceramic matrix can be made, for example, from aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), and silicon carbide (SiC), or other suitable ceramic materials. 
       FIG. 3B  shows additional details of airfoil section  68 . Respective inner surfaces  114 A,  114 B of suction-side CMC skin portion  98 A and pressure-side CMC skin portion  98 B can be bonded to outer surface(s)  112  of ceramic core  96  by way of corresponding suction- and pressure-side PTLP bonds  100 ,  102 . Suction-side CMC skin portion  98 A can be secured directly to an outer surface of ceramic core  96  via contiguous suction-side PTLP bond  100 , while pressure-side CMC skin portion  98 B can be secured indirectly to ceramic core  96  via a plurality of individual pressure-side PTLP bonds  102 . 
     PTLP bonds  100 ,  102  can include an alloyed interlayer having a melting temperature higher than a melting temperature of constituent elements defining the alloyed interlayer. The melting temperature is also higher than the bonding temperature. This results in high-temperature interlayer links between ceramic core  96  and CMC skin portions  98 A,  98 B which are more resilient and require less bonding area than a sintered connection between the ceramics. It also allows for the use of different ceramics and tailoring of mechanical and thermal properties of materials for core  96  and CMC skin portions  98 A,  98 B with much less concern for differential thermal expansion. 
       FIGS. 4A and 4B  show formation of PTLP bond  100  directly between inner surface  114 A of suction-side CMC skin portion  98 A and outer surface  112  of ceramic core  96 . A PTLP bond is one which has several similarities to brazed and diffusion-bonded connections, but which is formed at lower bonding temperatures than brazing and lower bonding pressures than diffusion bonding. Properly designed PTLP bonds can reduce intermaterial stresses and provide controlled diffusion between the different material interfaces. The lower temperatures of PTLP bond formation also mitigate potential microstructural weakening associated with other joining techniques. The resulting bond strength of alloyed interlayer  128  can be comparable to that of brazed, sintered, or diffusion-bonded connections and substantially maintains the structural integrity and composition of the substrates. 
       FIG. 4A  shows a precursor to PTLP bond  100 , PTLP bond assembly  120 , which includes refractory segment  122 , core-side foil layer  124 A, and skin-side foil layer  124 B. Foil layers  124 A,  124 B are shown as individual layers but one or both can alternatively comprise multiple foil layers. Refractory segment  122  can be, for example, nickel or an alloy thereof. Alternative refractory metals suitable for refractory segment  122  include gold, cobalt, copper, niobium, palladium, platinum, silicon, tantalum, titanium, vanadium, and alloys thereof. Foils  124 A,  124 B are selected so as to wet the ceramic substrate (here, ceramic core  96  and the ceramic matrix of CMC skin  98 A) at the bonding temperature. 
     As foil layers  124 A,  124 B are melted, thereby wetting the adjacent ceramic (i.e., core outer surface  112  and CMC skin inner surface  114 A), bond assembly  120  can then be maintained at a bonding temperature for a suitable time so as to homogenize the materials into PTLP bond  100  shown in  FIG. 4B  with alloyed interlayer  128 . 
       FIG. 5A  shows a configuration of PTLP bonding which incorporates thermal protection structure  104 . Thermal protection structure  104 , along with at least one PTLP bond  102 , is disposed between inner surface  114 B of pressure-side CMC skin portion  98 B and outer surface  112  of ceramic core  96 . 
     The configuration shown in  FIG. 5A  differs from  FIGS. 4A and 4B  in that a thermal protection structure is disposed across space  106  (shown in  FIG. 3B ) between surfaces  112 ,  114 B. One can take advantage of PTLP bonding to create a resilient high-melting-temperature and substantially uniform bond between two similar or dissimilar materials. With the configuration of  FIG. 5A , one can potentially utilize a third ceramic material for thermal protection structure  104 . The third material can be similar to the ceramic of one or both substrates. Alternatively, thermal protection structure  104  can be formed from a more thermally insulating ceramic relative to one or both ceramics of core  96  and CMC skin  98 B. 
     It can be seen that each of the plurality of thermal protection structures  104  (one shown in  FIG. 5A ) each have core side  132  and skin side  134  joined to a corresponding one of CMC skin inner surface  114 B and ceramic core outer surface  112 . Thermal protection structure  104  is shown here as an individual structure with both core side  132  and skin side  134  each joined to a corresponding one of ceramic core  96  and CMC skin  98 B by partial transient liquid phase (PTLP) bonds  102 . 
     PTLP bonds  102  can each be formed in a manner similar to that shown in  FIG. 4A , in which refractory segment  122  is sandwiched between at least one foil layer on either side to form a bond assembly  120 . Bond assemblies  120  are then heated to form PTLP bonds which have a higher melting temperature than the bonding temperature. This increased melting temperature is a result of isothermal solidification of alloyed interlayer  128  which mitigates the concern of remelting the bond. 
     Returning to  FIG. 5A , thermal protection structure  104  is shown as a separate structure bonded on either side to each substrate (core  96  and CMC skin  98 B). This is but one illustrative example configuration. It will also be appreciated that one or more portions of thermal protection structure  104  can be integrally formed into one or both of ceramic core  96  or skin  98 B. In one example, thermal protection structure  104  is integrally formed to ceramic core  96 , eliminating the need for one of PTLP bonds  102 . 
     In another example, shown in  FIG. 5B , interlocking or alternating thermal protection structures  104  can be formed on surfaces  112 ,  114 B.  FIG. 5B  shows a first thermal protection element  130 A and a second element  130 B joined by PTLP bond  132  to form alternate thermal protection structure  128 . A combination of such elements could also allow for appropriate mistake proofing by ensuring that the proper elements  130 A,  130 B line up for each thermal protection structure  128 . 
     Thermal protection structures  104 ,  128  (shown respectively in  FIGS. 5A and 5B ) can have any suitable cross-sectional geometry. In these examples, thermal protection structures  104 ,  128  can be an array of round or square projections. These and other example geometries are shown in commonly assigned U.S. patent application Ser. No. entitled: “Method For Joining Dissimilar Engine Components”, filed on an even date herewith. 
       FIG. 6  is a chart showing steps of method  200  for making a hybrid airfoiled component such as is shown in  FIGS. 2-5 . 
     Method  200  begins with step  202  of providing a ceramic airfoil core. This core may have a similar geometry as ceramic core  96  in the example above. However, other configurations are also possible, and is one benefit to the hybrid ceramic/CMC configuration. As noted in the preceding examples, the hybrid configuration allows for numerous complex shapes that would be too expensive or difficult to form out of a purely CMC airfoil. It also permits portions of the ceramic core to form leading and/or trailing edges of the airfoil to further simplify formation of the blade. 
     The ceramic airfoil core can be cast or otherwise formed out of a ceramic compound selected from one of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), silicon carbide (SiC), tungsten carbide (WC), and zirconium oxide (ZrO 2 ). 
     Step  204  includes placing a ceramic matrix composite (CMC) airfoil skin over at least a portion of the ceramic airfoil core. This can include placing one or more sheets of CMC material over the ceramic core such that they form an airfoil surface. The CMC skin can include a plurality of fibers selected from one or more of: silicon carbide (SiC), titanium carbide (TiC), aluminum oxide (Al 2 O 3 ), and carbon (C); and a ceramic matrix selected from one or more of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), and silicon carbide (SiC). 
     Step  206  can include, for example, placing a first thin metallic layer adjacent a core-side bonding surface, placing a second thin metallic layer on a skin-side bonding surface, and/or placing a refractory segment between the first and second thin metallic layers to form a bond assembly. Depending on the configuration of the desired airfoil, step  204  can be performed, in total or in part, after one or more of steps  206 ,  208 , and  210 . At least some of the constituents of the TLP and/or PTLP bond assembly can be positioned so as to prepare for steps  204 ,  208 , and/or  210 . 
     Optional step  208  involves spacing at least a portion of the CMC skin from the ceramic airfoil core. This can be done, for example, by providing a plurality of thermal protection structures between an outer surface of the ceramic airfoil core and an inner surface of the CMC airfoil skin. Each thermal protection structure can be provided a core side and a skin side joined to a corresponding one of the inner surface of the CMC airfoil skin and the outer surface of the ceramic airfoil core. Alternatively, the plurality of thermal protection structures can be integral with at least one of the inner surface of CMC airfoil skin and the outer surface of the ceramic airfoil core. 
     And at step  210 , the CMC skin is joined to the ceramic core to define an airfoil shape. As shown in  FIGS. 4A-5B , the CMC skin can be joined to the core at least in part by forming at least one of a transient liquid phase (TLP) and a partial transient liquid phase (PTLP) bond between the ceramic core and the CMC skin. The bond assembly is then heated to a bonding temperature to form the at least one bond which has an alloyed interlayer with a melting temperature higher than the bonding temperature. 
     As was shown in  FIG. 5B , the plurality of thermal protection structures can include at least one pair of opposed thermal protection elements, each of which includes a first structure projecting from the inner surface of the CMC airfoil skin, and a second structure projecting from the outer surface of the ceramic airfoil core. In these embodiments, joining step  206  can therefore include forming at least one partial transient liquid phase (PTLP) bond between each of the plurality of thermal protection structures and at least one of the ceramic airfoil core and the CMC airfoil skin. 
     DISCUSSION OF POSSIBLE EMBODIMENTS 
     The following are non-exclusive descriptions of possible embodiments of the present invention. 
     An airfoil comprises a core having a first surface, a skin having a second surface disposed over at least a portion of the first surface of the core, and at least one of a transient liquid phase (TLP) bond and a partial transient liquid phase (PTLP) bond. The at least one bond is disposed between the first surface and the second surface, joining the skin to the core. 
     The airfoil of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components: 
     An airfoil according to an exemplary embodiment of this disclosure, among other possible things includes a core having a first surface; a skin having a second surface disposed over at least a portion of the first surface of the core; and at least one of a transient liquid phase (TLP) bond and a partial transient liquid phase (PTLP) bond disposed between the first surface and the second surface, the bond joining the skin to the core. 
     A further embodiment of the foregoing airfoil, wherein the core comprises a ceramic compound selected from the group consisting of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), silicon carbide (SiC), tungsten carbide (WC), and zirconium oxide (ZrO 2 ). A further embodiment of any of the foregoing airfoils, wherein the core is monolithic. 
     A further embodiment of any of the foregoing airfoils, wherein the core defines at least one of: a leading edge of the airfoil, and a trailing edge of the airfoil. 
     A further embodiment of any of the foregoing airfoils, wherein the skin comprises at least one ceramic matrix composite (CMC) material. 
     A further embodiment of any of the foregoing airfoils, wherein the at least one CMC material comprises a plurality of ceramic fibers selected from one or more of: silicon carbide (SiC), titanium carbide (TiC), aluminum oxide (Al 2 O 3 ), and carbon (C). 
     A further embodiment of any of the foregoing airfoils, wherein the at least one CMC material comprises a ceramic matrix selected from one or more of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), and silicon carbide (SiC). 
     A further embodiment of any of the foregoing airfoils, wherein the skin is generally spaced from the core except proximate a location of the at least one bond. 
     A further embodiment of any of the foregoing airfoils, wherein the skin is generally spaced from the core by a plurality of thermal protection structures disposed therebetween, the plurality of thermal protection structures each having a core side and a skin side joined to corresponding one of the skin inner surface and the core outer surface. 
     A further embodiment of any of the foregoing airfoils, wherein at least one of the core side and the skin side is joined to the corresponding one of the CMC skin and the ceramic core by the at least one bond. 
     A further embodiment of any of the foregoing airfoils, wherein the at least one bond includes a PTLP bond comprising an alloyed interlayer having a melting temperature higher than a melting temperature of at least one constituent element defining the alloyed interlayer. 
     A further embodiment of any of the foregoing airfoils, wherein the skin includes at least one of a pressure-side sheet and a suction-side sheet. 
     A further embodiment of any of the foregoing airfoils, wherein the skin extends over the core proximate to at least one of a leading-edge portion of the core and a trailing-edge portion of the core. 
     A method for making a hybrid airfoiled component comprises providing a ceramic airfoil core. A ceramic matrix composite (CMC) airfoil skin is placed over at least a portion of the ceramic airfoil core. The CMC skin is joined to the ceramic core to define an airfoil shape. The joining step is performed at least in part by forming a partial transient liquid phase (PTLP) bond between the ceramic core and the CMC skin. 
     The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components: 
     A method for making a hybrid airfoil according to an exemplary embodiment of this disclosure, among other possible things includes: providing a ceramic airfoil core; placing a ceramic matrix composite (CMC) airfoil skin over at least a portion of the ceramic airfoil core; positioning at least one constituent element of a partial transient liquid phase (PTLP) bond assembly between the CMC skin to the ceramic core; and joining the CMC skin to the ceramic airfoil core, the joining step performed at least in part by forming a PTLP bond between the ceramic core and the CMC skin. 
     A further embodiment of the foregoing method, wherein the ceramic airfoil core comprises a ceramic compound selected from the group consisting of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), silicon carbide (SiC), tungsten carbide (WC), and zirconium oxide (ZrO 2 ). 
     A further embodiment of any of the foregoing methods, wherein the CMC skin comprises a plurality of fibers selected from the group consisting of: silicon carbide (SiC), titanium carbide (TiC), aluminum oxide (Al 2 O 3 ), and carbon (C); and a ceramic matrix selected from the group consisting of: aluminum oxide (Al 2 O 3 ), silicon nitride (Si 3 N 4 ), and silicon carbide (SiC). 
     A further embodiment of any of the foregoing methods, further comprising: spacing at least a portion of the CMC skin from the ceramic airfoil core. 
     A further embodiment of any of the foregoing methods, wherein spacing at least a portion of the CMC skin comprises: providing a plurality of thermal protection structures between an outer surface of the ceramic airfoil core and an inner surface of the CMC airfoil skin, the plurality of thermal protection structures each having a core side and a skin side joined to a corresponding one of the inner surface of the CMC airfoil skin and the outer surface of the ceramic airfoil core. 
     A further embodiment of any of the foregoing methods, wherein the plurality of thermal protection structures are integral with at least one of the inner surface of CMC airfoil skin and the outer surface of the ceramic airfoil core. 
     A further embodiment of any of the foregoing methods, wherein the plurality of thermal protection structures comprises at least one pair of opposed thermal protection structures, the pair of opposed thermal protection structures including a first structure projecting from the inner surface of the CMC airfoil skin, and a second structure projecting from the outer surface of the ceramic airfoil core. 
     A further embodiment of any of the foregoing methods, wherein the joining step comprises: forming at least one partial transient liquid phase (PTLP) bond between each of the plurality of thermal protection structures and at least one of: the ceramic airfoil core and the CMC airfoil skin. 
     A further embodiment of any of the foregoing methods, wherein the at least one constituent element of the PTLP bond assembly is selected from the group consisting of: placing a first thin metallic layer adjacent a core side bonding surface; placing a second thin metallic layer on a skin side bonding surface; and placing a refractory bond core between the first and second thin metallic layers to form a bond assembly. 
     A further embodiment of any of the foregoing methods, wherein the joining step comprises: heating the bond assembly to a bonding temperature to form the at least one PTLP bond, the at least one PTLP bond including an alloyed interlayer having a melting temperature higher than the bonding temperature. 
     A further embodiment of any of the foregoing methods, wherein the CMC skin defines at least a suction sidewall and a pressure sidewall of the airfoil shape. 
     A further embodiment of any of the foregoing methods, wherein the ceramic core defines at least one of: a leading edge of the airfoil, and a trailing edge of the airfoil. 
     Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.