Patent Application
20100251721
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.
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'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.
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.
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.
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
Referring to
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
As shown in
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 720fiber 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
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
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 720fiber 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-N550made 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
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
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
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
For example, A-N720CMC 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 610alumina, with a higher CTE than NEXTEL 720mullite 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.
