Abstract:
An assembly and method for controlling thermal stresses within ceramic-based articles when subjected to high temperatures while supported by a metallic article. The assembly includes a first body formed of a metallic material and having oppositely-disposed first and second surfaces, and a second body formed of a ceramic-based material and supported by the first body from the first surface thereof. The first and second bodies are located in a hot gas path such that the second body and the first surface of the first body are directly impinged by flowing hot gases. The assembly further includes a substantially uniform pattern of fins protruding from the second surface of the first body, and/or an interface structure between the first and second bodies that positively retains the second body to the first body and thermally insulates the first body from the first body.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates to ceramic matrix composite (CMC) articles, such as CMC components of gas turbine engines. More particularly, this invention is directed to an assembly and method for controlling thermally-induced stresses that exist within CMC articles when subjected to high temperatures while supported by metallic structures.  
         [0002]     Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of iron, nickel and cobalt-base superalloys. Nonetheless, superalloy components must often be air-cooled and/or protected by some form of thermal and/or environmental coating system in order to exhibit suitable service lives in certain sections of a gas turbine engine, such as the turbine, combustor and augmentor.  
         [0003]     As an example,  FIGS. 1 and 2  represent a nozzle segment  10  that is one of a number of nozzle segments that when connected together form an annular-shaped high pressure turbine (HPT) nozzle assembly of a gas turbine engine. The segment  10  is made up of multiple vanes  12 , each defining an airfoil and extending between outer and inner platforms (bands)  14  and  16 . The vanes  12  and platforms  14  and  16  can be formed separately and then assembled, such as by brazing the ends of each vane  12  within openings defined in the platforms  14  and  16 . Alternatively, the entire segment  10  can be formed as an integral casting. As a result of being located in the high pressure turbine section of the engine, the vanes  12  and the surfaces of the platforms  14  and  16  facing the vanes  12  are subjected to the hot combustion gases from the engine&#39;s combustor. Compressor bleed air is often supplied to the vanes  12  and platforms  14  and  16  to provide forced air cooling, such as with impingement and film cooling techniques. A thermal barrier coating (TBC) system is typically applied to the surfaces of the vanes  12  and platforms  14  and  16  exposed to the hot combustion gases to provide environmental protection and reduce heat transfer to the segment  10 .  
         [0004]     As one would expect, the nozzle segment  10  expands and contracts when heated and cooled, respectively, during transient engine operating conditions. Though provided with thermal protection to minimize service temperatures, the placement of the vanes  12  directly in the hot gas path results in the vanes  12  sustaining temperatures that are significantly higher than that experienced by the platforms  14  and  16 . As a result, in the situation where the vanes  12  and platforms  14  and  16  are formed of similar materials, such as nickel-base superalloys widely used, the vanes  12  will expand and contract more than the platforms  14  and  16 , thereby inducing significant thermally-induced strains and stresses in the segment  10 . Furthermore, significant temperature gradients will exist along the length of each vane  12  as a result of direct heating by the combustion gases and the platforms  14  and  16  behaving as heat sinks, thereby further increasing the thermally-induced stresses within the vanes  12 .  
         [0005]     In view of the higher operating temperatures sought for gas turbine engines to increase their efficiency, ceramic matrix composite (CMC) materials, such as silicon carbide (SiC)-containing CMC&#39;s, have been proposed as materials for certain components of gas turbine engines, including turbine vanes, blades, shrouds, combustor liners, and other high-temperature components of gas turbine engines. CMC materials have much higher temperature capabilities than the superalloys currently in use, and therefore better capable of withstanding the temperatures sustained in the hot gas path of a gas turbine engine. Depending on the exact materials used, CMC vanes  12  also typically have relatively lower coefficients of thermal expansion (CTE), resulting in lower strains and stresses thermally induced by the higher temperatures sustained by the vanes  12  as compared to the platforms  14  and  16 . However, CMC materials are also much less ductile and have significantly lower thermal conductivities than those of superalloys, rendering CMC vanes particularly prone to damage from thermally-induced stresses and strains. As a result, the general approach to implementing CMC vanes (e.g.,  12 ) supported with superalloy platforms (e.g.,  14  and  16 ) is to support the ends of the vanes with the platforms in a manner that allows differential thermal movement of the vanes relative to the platforms. Notable approaches are described in commonly-assigned U.S. Pat. No. 5,630,700 to Olsen et al. and U.S. Pat. No. 6,464,456 to Darolia et al.  
         [0006]     Notwithstanding the above advancements proposed by Olsen et al. and Darolia et al., further improvements would be desirable in order to fully implement nozzle assemblies having CMC vanes supported by metallic platforms. In particular, it would be desirable to improve the interface and attachment of the CMC vanes to the metallic platforms in order to not only reduce temperatures and thermally-induced stresses and strains, but also minimize the significant temperature gradients that can exist within a CMC vane  12 , which have been observed at times to exceed about 600° F. per inch (about 124° C./cm).  
       BRIEF SUMMARY OF THE INVENTION  
       [0007]     The present invention generally provides an assembly and method for controlling thermal stresses within a ceramic-based article when subjected to high temperatures while structurally supported by a metallic article.  
         [0008]     The assembly includes a first body formed of a metallic material and having oppositely-disposed first and second surfaces, and a second body formed of a ceramic-based material and supported by the first body from the first surface thereof. The first and second bodies are located in a hot gas path such that the second body and the first surface of the first body are directly impinged by flowing hot gases. According to a first aspect of the invention, the assembly further includes a substantially uniform pattern of fins protruding from the second surface of the first body. The fins are of sufficient size to increase the rigidity of the first body and promote heat transfer from the first body. According to a preferred aspect of the invention, the fins further serve to achieve more uniform temperatures within the first and second bodies and increase the stiffness of the first body to the extent that the thermal mass of the first body can be minimized.  
         [0009]     According to a second aspect of the invention, the assembly further has an interface structure between the first and second bodies. The interface structure comprises a resilient sealing member disposed between the first and second bodies and received in a recess in the first surface of the first body, and a ceramic saddle formed separately from the first and second bodies, received in the recess with the resilient sealing member, and disposed between the resilient sealing member and the second body. According to this aspect of the invention, the ceramic saddle and resilient sealing member cooperate to assist in positively retaining the second body to the first body and thermally insulating the first body from the first body in a manner that reduces thermal stresses within the second body, such as those caused by thermal gradients within the second body.  
         [0010]     The method of this invention generally entails forming a first body of a metallic material to have a first surface and a substantially uniform pattern of fins protruding from an oppositely-disposed second surface. The fins are of sufficient size to increase the rigidity of the first body and promote heat transfer from the first body. A second body of a ceramic-based material is then supported from the first surface of the first body. The resulting assembly is then placed in a hot gas path such that the second body and the first surface of the first body are directly impinged by flowing hot gases.  
         [0011]     It can be appreciated from the above that the invention is particularly applicable to nozzle assemblies of gas turbine engines, specifically where a ceramic-based vane is supported by and between a pair of platforms.  
         [0012]     Other objects and advantages of this invention will be better appreciated from the following detailed description.  
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0013]      FIGS. 1 and 2  are perspective and cross-sectional views, respectively, of a nozzle segment for a gas turbine engine and of a type that can benefit from the present invention.  
         [0014]      FIG. 3  is a perspective view showing a portion of a nozzle assembly configured in accordance with a preferred embodiment of the present invention.  
         [0015]      FIG. 4  is a cross-sectional view of the nozzle assembly of  FIG. 3  taken along section line  4 - 4  in  FIG. 3 .  
         [0016]      FIG. 5  is a sectional view of the nozzle assembly of  FIG. 3  taken along section line  5 - 5  in  FIG. 4 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0017]      FIGS. 3 through 5  schematically represent a portion of a nozzle segment  20  that, when assembled with similarly configured segments, forms an annular-shaped HPT nozzle assembly of a gas turbine engine. The nozzle segment  20  is depicted as including a single vane  22  supported by a single platform  24 , though it will be understood that multiple vanes can be supported by the platform  24  in combination with a second platform, resulting in a construction generally similar to that shown in  FIGS. 1 and 2 . The vane  22  is depicted as being hollow, though a variety of configurations are possible, including vanes configured to have struts, spars, inserts for mechanical support, baffles for enhanced internal cooling, etc. While the invention will be described in reference to a HPT nozzle assembly, it will be appreciated that the benefits of the invention can be applied to a variety of other components, including but not limited to low pressure turbine (LPT) nozzle assemblies and other hot section components of gas turbine engines.  
         [0018]     The platform  24  can be formed of such conventional materials as a single-crystal nickel, cobalt, or iron-base superalloy of a type suitable for use in gas turbine engines. Conventional practice has been to also form the vane  22  of the same or similar superalloy, such that the vane  22  and platform  24  would have similar CTE&#39;s and thermal conductivities to minimize thermally-induced strains and stresses during engine operation. However, according to the invention the vane  22  is formed of a ceramic-based material, more preferably a CMC material such as a SiC/SiC (reinforcement/matrix) CMC. However, it should be understood that this invention is applicable to the use of a variety of ceramic-based and metallic materials, as well as intermetallic materials such as nickel aluminides (NiAl), and particularly combinations of these materials that result in combinations with significantly different CTE&#39;s and/or thermal conductivities. For example, CMC materials of particular interest for the vane  22  may have CTE&#39;s and thermal conductivities in ranges of about 1.9×10 −6  to about 2.3×10 −6  in/in·° F. (about 8.7×10 −5  to about 1.9×10 −4  mm/mm·° C.) and about 7.8 to about 19.6 BTU/hr·ft·° F. (about 13.5 to about 33.9 W/mK), respectively, as compared to nickel-base superalloys whose CTE&#39;s and thermal conductivities of generally about 7.3×10 −6  to about 8.5×10 −6  in/in·° F. (about 3.3×10 −4  to about 3.9×10 −4  mm/mm·° C.) and about 6.8 to about 14.6 BTU/hr·ft·° F. (about 11.8 to about 25.3 W/mK), respectively. Such differences in CTE&#39;s can cause considerable differential thermal movement between the vane  22  and platform  24 , particularly during transient engine conditions when thermal conductivity and thermal mass also come into play.  
         [0019]     From  FIGS. 3 through 5 , the vane  22  is seen as being supported from a surface  26  of the platform  24 . In the particular embodiment shown in  FIGS. 3 through 5 , the platform  24  would be oriented radially inward from the vane  22  within the engine, and therefore may be referred to as the inner platform (or band) of the nozzle segment  20 . When installed on a gas turbine engine, the vane  22  and surface  26  of the platform  24  are directly impinged by hot combustion gases discharged by the combustor (not shown) and flowing along the hot gas path of the engine. As such, the vane  22  and platform  24  are both subjected to intense heating during engine operation. Consequently, a bleed air system may be employed that draws a portion of the compressed air from the engine&#39;s compressor (not shown) to cool the vane  22  and platform  24 , such as through backside cooling of the platform  24  by directing bleed air at the inner surface  28  of the platform  24  opposite the vane  22 , and/or by flowing bleed air through the vane  22 , a portion of which may be optionally discharged through film cooling holes (not shown) on the surface of the vane  22 . Such cooling techniques are well known in the art, and therefore do not require further explanation.  
         [0020]     As evident from  FIGS. 3 through 5 , the platform  24  is configured to include fins  30  protruding from its inner surface  28 . The fins  30  are of a sufficient size to serve as stiffeners that increase the rigidity of the platform  24 , thereby allowing the cross-sectional thickness of the platform  24  to be minimized to reduce the thermal mass of the platform  24 . As a result, the thermal inertia of the platform  24  is reduced, promoting more rapid heat transfer from the platform  24  with backside cooling. The fins  30  also preferably serve a secondary role of promoting radiation heat transfer from the platform  24  as a result of increased surface area from which heat can be radiated. Suitable dimensions for the fins  30  along the length of the platform  24  generally include a thickness (parallel to the surface  28 ) of about 2 to about 3 mm, and a height (normal to the surface  28 ) of about 2.5 to about 10 mm. In the embodiment depicted in  FIGS. 3 through 5 , the heights of the fins  30  increase immediately below the vane  22  to structurally accommodate a recess  38  (described in greater detail below) defined in the outer surface  26  of the platform  24 . To promote a more uniform temperature within the platform  24 , the fins  30  are preferably configured to define a substantially uniform pattern, such as the parallel pattern shown in  FIG. 3  through  5 , with a suitable uniform spacing between fins  30  of about 6 to about 13 mm. Because of the increased stiffness contributed by the fins  30  to the platform  24 , the cross-sectional thickness of the platform  24  (excluding the fins  30 ) can be reduced by, for example, about 15 to about 25 percent while maintaining the same level of stiffness, and simultaneously resulting in a thermal mass reduction of about 10 to 20% or more for the platform  24 . The fins  30  preferably extend the full circumferential length of the platform  24 , and are integrally formed with the remainder of the platform  24  such as during a casting process of any type known and used to produce platforms for gas turbine engine nozzle assemblies. Alternatively, the fins  30  could be formed separately and attached by welding, brazing, etc.  
         [0021]     The vane  22  is shown in  FIGS. 3 through 5  as mounted to the platform  24  with an interface structure  32  that provides a resilient, low thermal conductivity path between the vane  22  and platform  24 . The interface structure  32  is represented as including a seal  34  and saddle  36 , both of which are shown as being nested in the aforementioned recess  38  defined in the outer surface  26  of the platform  24 . The recess  38  provides positive axial and tangential retention of the vane  22 , the effect of which may be promoted by forming the recess  38  to extend through the inner surface  28  of the platform  24  and into the taller fins  30  shown immediately below the recess  38  in  FIG. 4 . Together the seal  34  and saddle  36  are shown in  FIG. 4  as completely filling the recess  38  and continuous between the opposing surfaces of the vane  22  and platform  24 .  
         [0022]     The seal  34  primarily provides the desired resilient interconnection between the vane  22  and platform  24 , while also serving to inhibit gas leakage and heat transfer between the vane  22  and the platform  24 . The seal  34  is preferably in the form of what may be termed a cloth seal, meaning a fabric-type sheet material woven from fibers. To withstand the high temperatures of the combustion gases within a high pressure turbine, the fibers are preferably formed of an oxide dispersion strengthened (ODS) material, though the use of other high-temperature materials is foreseeable. To have sufficient resilience and provide a desired level of thermal insulation, the seal  34  is preferably at least 2 to 3 mm thick (normal to the surface  26 ) and a porosity of about 0.5 to about 1.0%. Furthermore, the seal  34  is preferably continuous beneath the vane  22 , in contrast to the use of annular-shaped rope seals that surround the base of CMC vanes as proposed in the past. The seal  34  must also be sufficiently strong and stiff to resist compaction when under a compressive load between the vane  22  and platform  24 . In view of the above considerations, an example of a material suitable for use as the seal  34  is an ODS FeCrAl alloy commercially available from Plansee GmbH under the name PM2000.  
         [0023]     The saddle  36  is preferably formed of a ceramic-based material, more preferably a precast monolithic ceramic material such as SiC. A precast monolithic is believed to be preferred over a CMC material because of the desire for relatively precise control of the geometry of the saddle  36 , such as small radii fillets joining the portions of the saddle  36  parallel to and normal to the surface  26  of the platform  24 . Without interfering with the resilient connection provided by the seal  34 , the saddle  36  provides for positive retention of the vane  22  to the platform  24  by abutting a stepped shoulder  42  defined by the recess  38 . The abutting-supporting arrangement between the edge of the saddle  36  and the shoulder  42 , in combination with appropriate support at the end of the vane  22  opposite the platform  24 , also inhibits compaction of the seal  34  by the saddle  36 . In the preferred embodiment, the saddle  36  does not intentionally compress the seal  34 . The ceramic material of the saddle  36  also provides additional thermal insulation within the interface structure  32  to inhibit heat transfer between the vane  22  and the platform  24 . To have sufficient thickness for structural strength, the portions of the saddle  36  parallel to and normal to the surface  26  of the platform  24  are each preferably at least 2.5 to 5 mm thick. The depth for the stepped shoulder  42  below the surface  26  of the platform  24  is preferably equal to the thickness of the portion of the saddle  36  within the recess  38  so that that portion of the saddle  36  is generally flush with the surface  26 .  
         [0024]     The saddle  36 , vane  22 , and recess  38  in the surface  26  of the platform  24  are shown as having complementary configurations that form shiplap joints therebetween, as evident from  FIG. 4 . In particular, both the seal  34  and saddle  36  is depicted as having L-shaped cross-sections that nest with each other and with a recess  40  defined in a wall of the vane  22 , defining overlaps in both the plane parallel to the surface  26  and the plane normal to the surface  26 . The presence of in-series shiplap joints serves to reduce gas leakage between the vane  22  and platform  24 .  
         [0025]     As noted above, the fins  30  serve to reduce the temperature of the platform  24  by promoting radiation heat transfer from the platform  24  and reducing the thermal mass of the platform  24 . With the combination of the seal  34  and saddle  36  represented in  FIGS. 3 through 5 , the interface structure  32  thermally insulates the vane  22  from the platform  24 , thereby reducing thermal gradients within the vane  22  that could cause structural damage. The interface structure  32  further enables the vane  22  to be secured to the platform  24  in a manner that allows the vane  22  to expand and contract relative to the platform  24  during temperature excursions with reduced thermal-induced strains and stresses within the vane  22  that could cause the vane  22  to fracture during engine operation. More particularly, the vane  22  is able to expand and contract both radially and laterally, the latter of which includes the circumferential and axial directions of the engine. The end of the vane  22  opposite the platform  24  can be secured with a second platform (corresponding to the outer platform  14  of  FIG. 1 ) in the same manner or optionally in a manner consistent with the prior art, including the use of more rigid attachment techniques. With either approach, the interface structure  32  can potentially provide a sufficiently resilient connection between the vane  22  and its platform  24  to avoid the prior practice of constructing nozzle assemblies from multiple nozzle segments such as that shown in  FIGS. 1 and 2 , and instead forming the inner platform  24  (as well as the outer platform) as a single continuous ring.  
         [0026]     While the invention has been described in terms of a particular embodiment, it is apparent that other forms could be adopted by one skilled in the art. For example, the vane  22 , platform  24 , fins  30 , and interface structure  32  could be configured differently from that shown in the Figures while still achieving one or more of the intended objects of the invention. Accordingly, the scope of the invention is to be limited only by the following claims.