Patent Publication Number: US-8528339-B2

Title: Stacked laminate gas turbine component

Description:
FIELD OF THE INVENTION 
     This invention is directed generally to ceramic articles, and more particularly to ceramic articles that may be used in a turbine system as a replacement for metal components. 
     BACKGROUND OF THE INVENTION 
     Conventional gas turbine engines operate at high temperatures and therefore, many of the systems within the engine are formed from metals capable of withstanding the high temperature environments. For example, gas turbine systems often include ring segments that are stationary gas turbine components located between stationary vane segments at the tip of a rotating turbine blade or airfoil. Ring segments are exposed to high temperatures and high velocity combustion gases and are typically made from metal. While the metal is capable of withstanding the operating temperatures, the metal is often cooled to enhance the usable life of the ring segments. Many current ring segment designs use a metal ring segment attached either directly to a metal casing or support structure or attached to metal isolation rings that are attached to the metal casing or support structure. More recently, firing and/or operating temperatures of turbine systems have increased to improve engine performance. As a result, the ring segments have required more and more cooling to prevent overheating and premature failure. Even with thermal barrier coatings, active cooling is still necessary. 
     Ceramic materials, such as ceramic matrix composites, have higher temperature capabilities than metal alloys and therefore, do not require the same amount of cooling, resulting in a cooling air savings. Prior art ring segments made from CMC materials rely on shell-type structures with hooks or similar attachment features for carrying internal pressure loads. U.S. Pat. No. 6,113,349 and U.S. Pat. No. 6,315,519 illustrate ring segments with C-shaped hook attachments. Conventional ceramic matrix components are formed from layers of woven fibers positioned in planes and layers substantially parallel to the inner sealing surface of the ring segments. For cooled components, internal pressurization would load these attachment hooks in such a way as to cause high interlaminar tensile stresses. Other out-of-plane features common in laminated structures, such as T-joints, are also subject to high interlaminar stresses when loaded. One of the limitations of laminated ceramic matrix composite (CMC) materials, whether oxide or non-oxide based, is that their strength properties are not generally uniform in all directions (e.g. the interlaminar tensile strength is generally less than about 5% of the in-plane strength). Nonuniform fiber perform compaction in complex shapes and anisotropic shrinkage of matrix and fibers results in delamination defects in small radius corners and tightly curved sections, further reducing the already-low interlaminar properties. A further limitation of shell-type CMC construction is that the through-thickness thermal conductivity is lower than the in-plane conductivity, particularly for oxide based CMC&#39;s. Many applications of CMC require cooling, preferably convective cooling on one side, removing heat by through-thickness conduction. 
     An alternative to shell like CMC structures is to orient the CMC limited laminated structures in a configuration so as to minimize the negative effects of anisotropy. In this configuration laminated structures are oriented so that the fiber ends are normal to the gas path surfaces thereby eliminating the concern of poor interlaminar properties. Such orientation is referred to as stacked laminated structures. Stacked laminate construction does however have some drawbacks. It results in higher raw material use and thus higher waste as compared to other construction methods. Intricate shaping of the component is possible using the stacked laminate construction but cutting to form the shape results in wasted ceramic fabric during the fabrication process. The contemporary cutting practices used in stacked laminate construction typically results in a component having a greater amount of total ceramic fiber content. Such wasted ceramic fiber during cutting and greater ceramic fiber contents in the components greatly increases the cost of turbine components made from stacked laminate construction. Due to the cost of the materials, there is often a trade-off between the cost of the component and the desired properties of the component, such as higher thermal conductivity or higher creep strength. 
     Thus, a need exists for construction methods and structures for laminated ceramic composite components having a lower cost. There is a further need for such components having improved properties, such as higher thermal conductivity or higher creep strength. In addition, a need exists for a ceramic article that may be used as a replacement material for metal parts in turbine systems to improve the efficiencies of the turbine systems. 
     SUMMARY OF THE INVENTION 
     The exemplary embodiments described herein are directed to a stacked laminate component that may be used as a replacement for one or more metal components used in a turbine engine. The stacked laminate component can achieve multiple effects in a single structure by combining materials and selectively positioning those materials in accordance with critical and non-critical areas of the component. Lower cost components can also be achieved through use of lower cost materials being layered with superior materials, where the superior materials are generally positioned in the critical areas of the component. 
     In one aspect, a gas turbine component exposed to a hot gas path of a gas turbine is provided comprising a body with a radially inner surface along the hot gas path and a radially outer surface. The body has a plurality of layers being generally orthogonal to the radially inner surface. The plurality of layers comprise at least a first layer formed from a first material and a second layer formed from a second material. The first material is a ceramic matrix composite. 
     In another aspect, a gas turbine component exposed to a hot gas path of a gas turbine is provided comprising a body formed by a process of stacking and laminating layers to define a radially inner surface along the hot gas path. The layers can be generally orthogonal to the radially inner surface. The layers may be at least a first layer of a first material and a second layer of a second material. At least the first material is a ceramic matrix composite. The second material can have at least one of a higher thermal conductivity or a higher creep strength than the first material. 
     In another aspect, a method of manufacturing a gas turbine component is provided comprising: providing at least a first material and a second material; stacking and laminating the first and second materials to define a body comprising layers; and cutting the body. The first material is a ceramic matrix composite. The second material has at least one of a higher thermal conductivity or a higher creep strength than the first material. The first and second materials are arranged in alternating layers along at least a portion of the body. The layers are substantially orthogonal to a radially inner surface of the body. 
     The second material can be a ceramic matrix composite. The first and second layers may be positioned in an alternating pattern along the body. The second layer can be recessed from the first layer along the radially inner surface. The component can further comprise a coating on the radially inner surface, with the first layer extending into the coating. The first layer may be recessed from the second layer along the radially outer surface. The second layer may extend into the coating. 
     The component can further comprise an overwrap that imparts a compressive preload on the body. The overwrap can be designed to utilize a combination of properties of thermal expansion and processing shrinkage to provide a compressive preload on the body. The overwrap may be a ceramic matrix composite. The overwrap can be formed from a material having either a higher, or neutral coefficient of thermal expansion than the plurality of layers. The second material may be a sapphire fiber felt or a mullite whisker felt. The first and second layers may be positioned in an alternating pattern along at least a portion of the body. The component can be a ring seal segment, an airfoil, a platform, a vane or a combustor heat shield. 
     These and other embodiments are described in more detail below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention. 
         FIG. 1  is a front view of a ceramic matrix composite stacked laminate gas turbine component according to an exemplary embodiment of the invention. 
         FIG. 2  is a side view of a portion of the component of  FIG. 1  showing an exemplary embodiment of the stacked laminate construction of the invention. 
         FIG. 3  is an enlarged cross-sectional view of portion A of the component of  FIG. 2  showing an exemplary embodiment of the stacked laminate construction of the invention without the fiber overwrap. 
         FIG. 4  is an enlarged cross-sectional view of portion A of the component of  FIG. 2  showing another exemplary embodiment of the stacked laminate construction of the invention without the fiber overwrap. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Embodiments of the invention are directed to a construction for a ceramic matrix composite (CMC) turbine engine component. Aspects of the invention will be explained in connection with a ring seal segment, but the detailed description is intended only as exemplary. Embodiments of the invention are shown in  FIGS. 1-4 , but the present invention is not limited to the illustrated structure or application. 
     Referring to  FIGS. 1 through 3 , a ceramic matrix component is shown and generally represented by reference numeral  10 . The exemplary embodiment describes by way of example a CMC stacked laminate gas turbine component as a ring seal segment  10  for the turbine section of the gas turbine. However, it should be understood that the present disclosure contemplates gas turbine components of stacked laminate construction for other sections of the turbine engine, such as, for example, vanes, airfoils, vane platforms, combustor heat shields and the like. 
     Ring seal segment  10  can be used as a replacement for one or more metal components used in a turbine engine. Ring seal segment  10  can be formed from a plurality of layers  100  and  200  that are oriented unconventionally. For example, and as shown more clearly in  FIG. 2 , the layers  100  and  200  can be positioned generally or substantially orthogonal to an inner sealing surface or hot gas path side  20  such that the layers are orthogonal to the hot gas path  35  of the gas turbine. Such a configuration of layers  100  and  200  allows use of hooks and other attachment features where the loading is resisted by the CMC in the strongest direction of the CMC. In addition, the weaker interlaminar bonds are oriented in the lowest load direction of the ring segment  10 . 
     As shown in  FIG. 1 , the ring seal segment  10  can include a first foot  26  positioned on a backside surface  28  at a first end  40 . The backside surface  28  can be generally opposite the inner sealing surface  20 . The first foot  26  can extend generally orthogonally from the backside surface  28  and can include an outer attachment section  30 . The ring seal segment can also include a second foot  34  positioned on the backside surface  28  at a second end  42 . The second foot  34  can extend generally orthogonally from the backside surface  28  and can include an outer attachment section  36 . Outer attachment sections  30  and  36  can be used for attachment to the gas turbine by an attachment structure (not shown). Such attachment structures are known in the art. 
     The ring seal segment  10  can include an abradable and/or insulative coating  50  on the inner sealing surface  20 . The coating  50  can be any conventional or not yet developed abradable and/or insulative coating. The coating  50  can be attached to the inner sealing surface  20  through any appropriate method, such as, for example, an intermediate adhesive layer or other bond-enhancing material, and can include insulative properties in some embodiments. The coating  50  can be, for example, a friable graded insulation (FGI). Various examples of FGI coatings are disclosed in U.S. Pat. Nos. 6,676,783; 6,670,046; 6,641,907; 6,287,511; 6,235,370; and 6,013,592. 
     The coating  50  can be applied over at least a portion of the inner sealing surface  20 . In one embodiment, the coating  50  can completely cover the inner sealing surface  20 . The thickness of the coating  50  can be substantially uniform, but, in some cases, it can be preferred if the thickness of the coating  50  is non-uniform. The variation in thickness of the coating  50  can occur in one or more directions, or it can vary in localized regions. 
     Layers  100  and  200  have differing properties that allow for selective control of the characteristics of the ring seal segment  10  in different portions of the segment. For example, layer  100  can be a CMC having high temperature tolerances and high strength such as NEXTEL 720 fiber reinforced alumina composite (A-N720) made by COI Ceramics Inc. Layer  200  can be a material having a higher thermal conductivity than layer  100 . For example, layer  200  can be a monolithic or CMC such as (A-N610) made by COIC Ceramics Inc. (A-N191) made by Saint Gobain, a ceria-based refractory or other relatively high thermally conductive materials. Such materials can be stacked with layer  100  to enhance the heat transfer from the hot gas path  35  to the backside surface or cool side  28  of the ring seal segment  10 . 
     To further increase the heat transfer surface area and the convection coefficient, the layers  200  can protrude or extend beyond the layers  100  along the cool side  28  as shown at ends  205  of the layers  200 . Layer  200  can also be recessed from the layers  100  along the hot gas path side or inner sealing surface  20  as shown at ends  210  of the layers  200 . By recessing ends  210 , layers  200  can be protected from the higher temperatures to which the inner sealing surface  20  is exposed. This is especially significant where materials are being used for layers  200  that have high thermal conductivity but only limited temperature tolerance. 
     To enhance the bond between the coating  50  and the inner sealing surface  20  of the ring seal segment  10 , the ends  105  of layers  100  can protrude into the coating. Such an arrangement provides greater surface area for adhesion of the coating  50  to the inner sealing surface  20 , with the added benefit of giving a mechanically interlocking feature that provides additional bonding benefits for the coating material  20 . 
     Portion A of  FIG. 2  shows layers  100  and  200  arranged in alternating columns. However, the present disclosure contemplates the use of other patterns of layers  100  and  200 . Ring seal segment  10  can also have more than two layers of different materials. The particular arrangement of layering can be chosen to focus the superior properties of the materials on those portions of the ring seal segment  10  or other gas turbine component that can take the most advantage of the properties. The exemplary embodiment has one layer that is made from a CMC material as described above. The additional layer or layers can be CMC or other such materials that allow for multiple effects through material properties to be achieved in a single structure. 
     For example, layers  200  having higher thermal conductivity can be arranged in an alternating pattern with layers  100  along the mid-section  15  of the ring seal segment  10 , while the adjacent ends  40  and  42  of the ring seal segment are composed only of layers  100 . Such a non-uniform arrangement of the layers  100  and  200  can increase the heat transfer along the mid-section  15  where the cool side  28  is in proximity to the hot gas path side  20  while maintaining strength along the ends  40  and  42  of the ring seal segment  10  that are in proximity to the attachment sections  30  and  36 . This results in lower average temperatures of layers  100  thereby improving the usable strength of this layer. 
     Ring seal segment  10  can have layers  100  and  200  of substantially equal thickness as shown in  FIG. 3 . However, the present disclosure contemplates the use of varying thicknesses of layers  100  and  200 . For example, layers  200  of increased thickness can be positioned along mid-section  15  to enhance heat transfer between the inner sealing surface  20  and the cool or backside surface  28 , while layers  200  of decreased thickness can be used along ends  40  and  42  of the ring seal segment  10  where there is less need for heat transfer. Similarly, the thickness of layer  100  or the thickness of any additional layers of materials that are utilized in ring seal segment  10  can be varied. An example of where layer  100  might need to be thicker would be at either end of the laminated structure which might typically be exposed to higher thermal stresses. 
     Layers  100  and  200  can also be chosen so as to make ring seal segment  10  more cost effective. For example, layer  100  can be a CMC having high temperature tolerances and strength such as NEXTEL 720 fiber reinforced alumina composite (A-N720) made by COI Ceramics Inc. Layer  200  can be a material having a lower cost than that of layer  100 . For example, layer  200  can be a monolithic or CMC such as AS-N550 made by COIC Ceramics Inc. (A-N191) made by Saint Gobain, FGI, ZIRCAR fiber board, a ceria-based refractory or other cost effective materials. Such materials can be stacked with layer  100  to reduce the overall cost of the ring seal segment  10 . Where the cost effective material has lower temperature tolerance, layers  200  can be protected from the higher temperatures to which the inner sealing surface  20  is exposed by being recessed from the layers  100  along the inner sealing surface as shown at ends  210  of the layers  200 . The pattern of layering of the cost effective material of layers  200  with respect to layers  100  can be chosen so as to position the layers  200  in the less critical areas of the ring seal segment  10  and position layers  100  in the more critical areas. The critical areas can include those areas of ring seal segment  10  that are exposed to higher temperatures and those areas that are exposed to higher stresses. Although, the present disclosure contemplates defining critical areas for positioning of the layers  100  based upon the particular superior properties of the material of layers  100 . 
     Referring to  FIG. 4 , ring seal segment  10  can have layers  300  and  400  with differing properties that allow for selective control of the properties in different portions of the ring seal segment  10 . For example, layer  300  can be a CMC having high temperature tolerances and strength such as the A-N720 described above. Layer  400  can be a material having higher creep deformation resistance than that of layer  300 . For example, layer  400  can be a sapphire fiber felt such as one made by Foster-Miller, a mullite whisker felt such as one made by NSWC, or other highly creep resistant materials. Additionally, layer  400  can have the same nominal composition as layer  300 , but processed to a higher temperature. Coarsening of grain structure by such higher temperature processing can reduce strength, but will impart improved creep resistance for layer  400 . Layer  400  can also be a continuous fiber CMC with additions of single crystal fibers of whiskers. Such materials of layer  400  can be stacked with layer  300  to mitigate against creep deformation. 
     To enhance the bond between the coating  50  and the inner sealing surface  20  of the ring seal segment  10 , the ends  410  of layers  400  can protrude into the coating. Such an arrangement provides greater surface area for adhesion of the coating  50  to the inner sealing surface  20 , as well as a mechanical lock of the stronger layers  400  with the coating. The ends  305  and  405  of the layers  300  and  400  can be flush with each other as shown in  FIG. 3  or can be offset. 
       FIG. 4  shows layers  300  and  400  arranged in alternating columns. However, the present disclosure contemplates the use of other patterns of layers  300  and  400 . This is especially significant where costly material, such as a sapphire fiber felt, is being used for layer  400 . For example, the sapphire fiber felt of layers  400  or any other creep resistant material, can be positioned along the critical portions of the ring seal segment  10  where creep deformation is at its highest and can be used sparingly, if at all, along those portions of the ring seal segment  10  where creep deformation is at its lowest. As described above, ring seal segment  10  can also have more than two layers of different materials. The particular arrangement of layering can be chosen to focus the superior properties of the materials on those portions of the ring seal segment  10  that can take the most advantage of the properties. Ring seal segment  10  can have layers  300  and  400  of substantially equal thickness. However, the present disclosure contemplates the use of varying thicknesses of layers  300  and  400 . Similarly, the thickness of layer  100  or the thickness of any additional layers of materials that are utilized can be varied. 
     The processing of layers  100 ,  200 ,  300  and/or  400  to form ring seal segment  10  can be any appropriate technique including co-processing, post-process bonding and any combination thereof. Cutting techniques such as water jet cutting and laser cutting can be used to form the final shape of the gas turbine component such as forming the ring seal segments  10  described above. 
     Other types of ceramic materials can be used for layers  100 ,  200 ,  300  and/or  400 , as well as any additional layers that are being utilized in the gas turbine component. Examples of such ceramic materials can include, but are not limited to, cerium oxide, alumina, zirconia, glass, silicon carbide, silicon nitride, sapphire, cordierite, mullite, magnesium oxide, zirconium oxide, boron carbide, aluminum oxide, tin oxide, scandium oxide, hafnium oxide, yttrium oxide, spinel, garnet, steatite, lava, aluminum nitride, iron oxide, aluminosilicate, porcelain, forsterite or combinations thereof, as well as any other crystalline inorganic nonmetallic material or clay. Other types of non-ceramic materials can also be used for layers  200  and/or  400 , as well as any additional layers that are being utilized in the gas turbine component. 
     The ring seal segment  10  can include the use of a strengthening mechanism  500  selected to provide reinforcement to the ring seal segment to increase the strength of the layers  100 ,  200 ,  300  and/or  400 , an example of which is shown in  FIG. 2 . The strengthening mechanism  500  can be selected such that it is located within one or more locations of the ceramic article. As such, the ring seal segment  10  or other gas turbine component structured in accordance with the exemplary embodiments, can be used as a replacement for one or more parts in a turbine system that are typically metal, thereby enabling the greater temperature capacity of the ceramic materials to be utilized such that the efficiencies of the turbine systems can be increased relative to prior art systems. 
     The strengthening mechanism  500  is selected to be positioned with respect to the ring seal segment  10  to help reinforce the segment and/or prevent delamination of the CMC layers that compose the segment. Therefore, the strengthening mechanism  500  serves to reinforce the layers  100 ,  200 ,  300  and/or  400 , especially normal to the plane of the layers and/or to help inhibit separation of the layers. The number, size, shape and location of the strengthening mechanisms  500  used can be optimized based upon one or more factors including, but not limited to, the local stresses to be applied to the ring seal segments  10 , the materials used for layers  100 ,  200 ,  300  and/or  400  and/or the type of strengthening mechanism  14 . 
     The strengthening mechanisms  500  can place the layers  100 ,  200 ,  300  and/or  400  under compression in a direction generally parallel to the inner sealing surface  20  of the ring seal segment  10 . In one embodiment, the strengthening mechanism  500  can be a CMC over-wrap that is wrapped around a portion of the ring seal segments  10 . The over-wrap  500  can be composed of a ceramic matrix composite material or other appropriate materials. As shown in  FIG. 2 , the over-wrap  500  can be in the form of a fiber, a sheet, a fabric, a tow, braided strips or other appropriate materials. A combination of different over-wraps  500  can also be used. The over-wrap  500  can be placed around the ceramic article in one or more locations to help reinforce the ring seal segment  10 . The over-wrap  500  can be placed around the ring seal segment  10  after formation of the ring seal segment or during processing or formation of the ring seal segment. In one embodiment, the over-wrap  500  is placed around the ring seal segment  10  after the ring seal segment is fully or nearly fully fired such that the natural shrinkage of the CMC over-wrap, such as during a secondary processing, can be used to induce residual compressive stress on the ring seal segment. 
     For example, A-N720 CMC can be used to form the over-wrap  500 . When the over-wrap  500  is placed onto the fully fired layers  100 ,  200 ,  300  and/or  400 , the over-wrap can result in a differential shrinkage strain of 0.1% to 0.3%, depending on the firing temperature of the final assembly. This strain can impose an interlaminar compressive stress on the laminate stack, thus adding to the load-carrying capability in this direction. The CMC over-wrap  500  can also be formed from a material having a higher coefficient of thermal expansion than layers  100 ,  200 ,  300  and/or  400 . In this embodiment, during secondary processing, the overwrap shrinks to compressively load the stacked laminate structure. During cool-down, the compressive load is relaxed and will eventually transform to a zero compressive load at room temperature. However, during operation, the stacked laminate structure is at a higher temperature than the overwrap. This temperature differential results in the overwrap maintaining a compressive load on the stacked laminate structure. 
     In addition, the CMC over-wrap  500  can be formed from a different composition with different sintering shrinkage than the layers  100 ,  200 ,  300  and/or  400 , such as a material with a greater sintering shrinkage. The process of coupling the over-wrap  500  to the layers  100 ,  200 ,  300  and/or  400  can include securing the layers together with at least one strengthening mechanism  500  and applying a processing temperature to the over-wrap to provide a defined shrinkage differential and compressive preload to the plurality of layers. The over-wrap  500  and the layers  100 ,  200 ,  300  and/or  400  can also be subjected to an intermediate firing stage before application of the over-wrap so that shrinkage can be controlled at final firing of the ring seal segment  10 . 
     In an alternative embodiment, alternative fibers can be used for the over-wrap material  500  to achieve further shrinkage and/or coefficient of thermal expansion (CTE) mismatch pre-stressing. For example, in the case above, if the overwrap fiber is NEXTEL 610 alumina, with a higher CTE than NEXTEL 720 mullite fiber, a differential shrinkage of 0.2% to 1.0% can be achieved by a combination of CTE and sintering shrinkage. In some embodiments, the over-wrap  500  can be located in, or adjacent to, regions of interlaminar tensile stress. For thermally induced stresses, it can be beneficial to locate the overwrap  500  around the neutral axis of bending. 
     In another embodiment, the over-wrap material  500  can be processed after placement on the ring seal segment  10 . This secondary processing can be used to permit for alternative CMC materials to be used for the over-wrap  500 , particularly if the over-wrap is to be located within a cooler region removed from the inner sealing surface  20  of the ring seal segment  10  when in use. For example, an aluminosilicate matrix material having superior bond strength and increased shrinkage can be used in the cooler regions of the over-wrap  500 . 
     The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.