Patent Description:
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature resistance. Ceramics, such as ceramic matrix composite ("CMC") materials, are also being considered for airfoils. CMCs have high temperature resistance. Despite this attribute, there are unique challenges to implementing CMCs in airfoils.

<CIT> discloses a ceramic matrix composite blade for use in a gas turbine engine is disclosed. The ceramic matrix composite blade includes a root, an airfoil, and a platform located between the root and the airfoil.

<CIT> discloses a rotor assembly comprising a rotary structure extending circumferentially about an axial centerline of a gas turbine engine and an airfoil having a root and a tip. The root is coupled to the rotary structure and has a bulbous shape. The airfoil is formed from a plurality of composite plies, a portion of which defines at the root first and second end surfaces, which are in contact with the rotary structure and together define a chisel-shaped end of the root.

<CIT> discloses hybrid turbine airfoil components containing ceramic material. The components comprise an airfoil portion, a nub, a platform portion therebetween, and a dovetail portion on the nub.

An airfoil according to a first aspect of the present invention includes a fiber-reinforced composite airfoil core that defines an airfoil portion and a root portion. The fiber-reinforced composite airfoil core is subject to core thermal growth. There is fiber-reinforced composite wrapping that wraps around the root portion. The fiber-reinforced composite wrapping is subject to wrapping thermal growth. There is a buffer layer between the root portion and the platform. The buffer layer absorbs a mismatch between the core thermal growth and the platform thermal growth.

In a further embodiment of any of the foregoing embodiments, the root portion has an axial face, and the buffer layer is on the axial face.

In a further embodiment of any of the foregoing embodiments, the root portion includes circumferential faces that substantially exclude the buffer layer.

In a further embodiment of any of the foregoing embodiments, the buffer layer wraps around edges of the axial face.

In a further embodiment of any of the foregoing embodiments, the buffer layer is a coating.

In a further embodiment of any of the foregoing embodiments, the coating is selected from the group consisting of alumina, silicon carbide, and combinations thereof.

In a further embodiment of any of the foregoing embodiments, the buffer layer includes hollow spheres disposed in a matrix.

In a further embodiment of any of the foregoing embodiments, the matrix is selected from the group consisting of alumina, silicon carbide, and combinations thereof.

According to a further aspect of the invention, a gas turbine engine is disclosed according to claim <NUM>.

<FIG> illustrates a representative airfoil <NUM>. In this example, the airfoil is from the turbine section of the engine <NUM> and is a rotatable blade that includes an airfoil core <NUM> and a ceramic matrix composite wrapping <NUM>. Although the examples below may be presented with respect to ceramic matrix composites, this disclosure can also be applied to other fiber-reinforced composite airfoils in other locations in the engine <NUM> which would benefit here from. For instance, the examples of this disclosure may be applied to airfoils in the compressor section <NUM> that utilize high-temperature polymers, such as fiber-reinforced composites having bismaleimide matrices.

In the illustrated example, the ceramic matrix composite wrapping <NUM> defines a platform <NUM> (integrated). Although the airfoil core <NUM> and the ceramic matrix composite wrapping <NUM> are integrated into a single component, the airfoil <NUM>, the airfoil core <NUM> and the ceramic matrix composite wrapping <NUM> are formed of distinct bodies. Alternatively, as shown in <FIG>, the ceramic matrix composite wrapping <NUM> excludes a platform and platforms <NUM> are provided as separate pieces. The examples herein are understood to refer to ceramic matrix composite wrapping <NUM> both with and without integrated platforms.

Referring to <FIG>, the airfoil core <NUM> defines several portions, including an airfoil portion <NUM> and a root portion <NUM>. The airfoil portion <NUM> has an aerodynamic profile, while the root portion <NUM> has a dovetail profile. The airfoil core <NUM> is formed of a ceramic matrix composite ("CMC") <NUM> (or bismaleimide fiber-reinforced composite as indicated above), which is shown in a cutaway portion in <FIG> and includes fibers 70a disposed in a ceramic matrix 70b (or bismaleimide). The fibers 70a may be provided in a fiber structure, represented at 70c, such as but not limited to unidirectional, woven, or braided structures. Example fibers 70a are silicon-containing ceramic fibers, such as silicon carbide (SiC) fibers or silicon nitride (Si<NUM>N<NUM>) fibers. Other types of ceramic fibers or carbon fibers may alternatively be used. Example ceramic matrices 70b are silicon-containing ceramics, such as silicon carbide (SiC) or silicon nitride (Si<NUM>N<NUM>). Other ceramics may alternatively be used.

In the illustrated example, the ceramic matrix composite wrapping <NUM> includes an endwall portion <NUM> and a root portion <NUM>. If there is no integrated platform <NUM>, the endwall portion <NUM> is excluded such that the ceramic matrix composite wrapping <NUM> includes only the root portion <NUM>. The root portion <NUM> surrounds the root portion <NUM> of the airfoil core <NUM> and also has a dovetail profile. The ceramic matrix composite wrapping <NUM> is also formed of a CMC <NUM> (or bismaleimide fiber-reinforced composite as indicated above), which is shown in a cutaway portion in <FIG> and includes fibers 76a disposed in a ceramic matrix 76b (or bismaleimide). The fibers 76a and the ceramic matrix 76b may be any of the materials described above for the CMC <NUM>. The fibers 76a may be provided in a fiber structure, represented at 76c, such as but not limited to unidirectional, woven, or braided structures. The fiber structure 76c of the ceramic matrix composite wrapping <NUM> wraps around the root portion <NUM> of the airfoil core <NUM>, substantially covering the axial and circumferential faces of the root portion <NUM> of the airfoil core <NUM>. Most typically, the CMC <NUM> of the ceramic matrix composite wrapping <NUM> and the CMC <NUM> of the airfoil core <NUM> will be of the same composition with regard to the chemistry and amounts of the fibers 70a/76a and the matrices 70b/76b. The CMCs <NUM>/<NUM> may or may not have the same fiber structure.

In wrapped configurations such as in the airfoil <NUM>, even relatively small thermal growth in a root portion of an airfoil core can cause relatively high stress in a wrapping, particularly at the axial ends due to axial growth. In this regard, as shown in the sectioned view of the airfoil <NUM> in <FIG> (see also sectioning line in <FIG>), the airfoil <NUM> includes a buffer layer <NUM> located between the root portion <NUM> of the airfoil core <NUM> and the root portion <NUM> of the ceramic matrix composite wrapping <NUM>. The buffer layer <NUM> absorbs relative thermal growth displacements between the root portions <NUM>/<NUM>, to facilitate mitigation of stress on the ceramic matrix composite wrapping <NUM> from differences in thermal growth between the root portions <NUM>/<NUM>. For example, the buffer layer <NUM> is compressible under the stress such that it becomes reduced in volume.

In general, the buffer layer <NUM> will be formed of ceramic material in order to avoid substantial damage during thermal processing of the CMCs or other fiber-reinforced composite. For instance, in the illustrated example, the buffer layer <NUM> is a coating. In examples, the coating is a ceramic coating such as alumina, silicon carbide, silicon nitride, silicate, oxide, boron carbide, or combinations thereof, or a metallic coating, such as elemental silicon. Other types of ceramic can additionally or alternatively be used, as long the buffer layer <NUM> is compressible relative to the thermal growth and stresses of the root portions <NUM>/<NUM>. As will be appreciated, the thermal growth and stresses will vary somewhat in accordance with the composition of the CMCs and their fiber structures, as well as the size and geometry of the airfoil <NUM>. In general, however, the root portions <NUM>/<NUM> will have a conforming dovetail shape and will be at least about <NUM> (millimeters) in axial length. The root portions <NUM>/<NUM> will typically not be more than about <NUM> in axial length. Thermal growth and stresses can be measured or estimated experimentally or by computer simulation for given materials and designs, and given this disclosure those skilled in the art will thus be readily able to identify useful buffer layers <NUM>.

In the illustrated example, the buffer layer <NUM> is located adjacent an axial face 68a of the root portion <NUM> of the airfoil core <NUM> in order to facilitate the absorbance of thermal growth in the axial direction. This is also shown in the isolated view of the airfoil core <NUM> in <FIG>, which excludes the ceramic matrix composite wrapping <NUM> in order to observe the buffer layer <NUM>. The axial face 68a may be the forward or aft axial face of the airfoil core <NUM>, or buffer layers <NUM> may be provided at both the forward and aft axial faces. In this example, the circumferential faces 68b/68c of the root portion <NUM> of the airfoil core <NUM> exclude the buffer layer <NUM>. The buffer layer <NUM> is primarily or exclusively on the axial face 68a, however, as shown in <FIG> the buffer layer <NUM> may alternatively wrap around the localized edges of the axial face 68a, which may also experience thermal stress.

The axial thickness of the buffer layer <NUM>, represented at 78a, is selected to be greater than the maximum thermal growth displacement between the root portions <NUM>/<NUM>. For example, the thickness 78a of the buffer layer <NUM> is from <NUM> millimeters to <NUM> millimeters.

<FIG> illustrates the same sectioned view as in <FIG> but under an elevated thermal state (temperature). For example, in the engine <NUM>, the airfoil portion (<NUM>) temperature may be from approximately <NUM> to <NUM>, whereas the root portions 68a/68b/68c may be from approximately <NUM> to <NUM>.

In the elevated thermal state there is a differential thermal expansion between the root portions <NUM>/<NUM> in the axial direction such that the buffer layer <NUM> is compressed between the root portions <NUM>/<NUM> and becomes reduced in volume. For instance, the buffer layer <NUM> may non-destructively compress or destructively compress (i.e., fracture). If the buffer layer <NUM> wraps around the localized edges of the axial face 68a (<FIG>), the "arms" of the buffer layer <NUM> that wrap around may also be compressed. There may also be other forces on the airfoil <NUM>, such as centrifugal forces from rotation.

The buffer layer <NUM> may be engineered in composition, structure, or both in order to provide a preset level of compressibility. For instance, as shown in the example in <FIG>, the buffer layer <NUM> is a ceramic coating that includes hollow spheres 80a disposed in a matrix 80b. For example, the matrix 80b is alumina, silicon carbide, silicate, oxide, or combinations thereof and the spheres 80a are glass spheres. The spheres 80a are relatively weak and provide porosity in the coating to facilitate compressibility.

The ceramic matrix composite wrapping <NUM> may also facilitate reinforcement of the root portion <NUM> of the airfoil core <NUM>. For example, the ceramic matrix composite wrapping <NUM> may serve to contain CMC plies in the root portion <NUM>. Fibers in the CMC of the ceramic matrix composite wrapping <NUM> run in the hoop direction, providing high stiffness in tension and thus containing the root portion <NUM> to enhance interlaminar strength between CMC plies in the root portion <NUM>.

<FIG> illustrates an example in which the airfoil <NUM> does not fall within the scope of the invention and includes a frangible layer <NUM>. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements. Here, the frangible layer <NUM> is similar to the buffer layer <NUM> but is designed to ensure fracture in compression under the elevated thermal state. As will be appreciated, since the frangible layer <NUM> is designed to fracture, the temperature at which the frangible layer <NUM> fractures corresponds to a typical use temperature of the airfoil <NUM>, such as the temperature at cruise. In one example, the frangible layer <NUM> is a porous ceramic coating of alumina, silicon carbide, silicate, oxide, or combinations thereof that has a compressive strength that is below the stress produced from differences in thermal growth between the root portions <NUM>/<NUM>. For example, the porosity, by volume percent, is from <NUM>% to <NUM>%. Thermal growth and stresses can be measured or estimated experimentally or by computer simulation for given materials and designs, and given this disclosure those skilled in the art will thus be readily able to identify useful frangible layers <NUM> and compressive strengths. Most typically, however, the compressive strength of the frangible layer <NUM> will be below <NUM> Mpa for relatively lower porosity compositions to less than <NUM> MPa for relatively highly porous frangible layers at a temperature of approximately <NUM>.

<FIG> illustrates the same sectioned view as in <FIG> but under an elevated thermal state as described above. The frangible layer <NUM> is compressed between the root portions <NUM>/<NUM>, reduces in volume, and fractures, i.e., is crushed, to absorb the relative thermal growth displacements between the root portions <NUM>/<NUM>. Once initially crushed, the frangible layer <NUM> may or may not be crushed further upon additional thermal growth or cycles of expansion and contraction. The space that the frangible layer <NUM> occupies serves as an expansion gap after fracture to accommodate the relative thermal growth displacements between the root portions <NUM>/<NUM>.

The buffer layer <NUM> or frangible layer <NUM> (collectively "layers <NUM>/<NUM>") also serve to facilitate fabrication of the airfoil <NUM>. For instance, during fabrication of the airfoil <NUM>, the airfoil core <NUM> is fully or partially formed, and the ceramic matrix composite wrapping <NUM> is then formed around the root portion <NUM> of the airfoil core <NUM>. But for the physical presence of the layer <NUM>/<NUM>, there may be considerable difficulty in wrapping the ceramic matrix composite wrapping <NUM> to the required geometry and dimensions. For instance, if the region where the layer <NUM>/<NUM> resides were instead left void as a hollow expansion gap, there would be no structure on which to wrap the ceramic matrix composite wrapping <NUM> to ensure that the platform <NUM> is properly located. Rather, in the airfoil <NUM>, the layers <NUM>/<NUM> also serve as a mandrel over which the ceramic matrix composite wrapping <NUM> is wrapped. The layers <NUM>/<NUM> support the ceramic matrix composite wrapping <NUM> and thus can ensure that the ceramic matrix composite wrapping 64is properly located. Once the ceramic matrix composite wrapping <NUM> is wrapped and rigidized or consolidated via the matrix 76b, the ceramic matrix composite wrapping <NUM> is self-supporting. There is then no further need after fabrication of the ceramic matrix composite wrapping <NUM> for the layer <NUM>/<NUM> to support the ceramic matrix composite wrapping <NUM> and, therefore, even if the layer <NUM>/<NUM> fractures, the fracturing does not debit the performance of the airfoil <NUM>.

In one further example fabrication, the airfoil core <NUM> is fully or substantially fully formed. For instance, the fiber structure 70c is provided, such as by a lay-up of plies that are formed into the desired geometry. The fiber structure 70c is then fully or partially densified with the matrix 70b. For instance, densification includes, but is not limited to, chemical vapor deposition of the matrix 70b. Subsequently, the layer <NUM>/<NUM> is arranged adjacent the axial face or faces 68a of the root portion <NUM> of the airfoil core <NUM>. For instance, the layer <NUM>/<NUM> is deposited onto the axial face 68a by a process such as but not limited to plasma spraying. Once the layer <NUM>/<NUM> is fully formed, the fiber structure 76c of the ceramic matrix composite wrapping <NUM> is wrapped around the root portion <NUM> of the airfoil core <NUM> and the layer <NUM>/<NUM>, such as by a lay-up of plies that are formed into the desired geometry. The fiber structure 76c is then densified with the matrix 76b, such as by chemical vapor deposition of the matrix 76b. If the fiber structure 70c of the airfoil core <NUM> was not fully densified, the fiber structure 70c may be further densified during densification of the fiber structure 76c. The layer <NUM>/<NUM> thus becomes trapped or sandwiched between the airfoil core <NUM> and the ceramic matrix composite wrapping <NUM>.

Claim 1:
An airfoil (<NUM>) comprising:
a fiber-reinforced composite airfoil core (<NUM>) defining an airfoil portion (<NUM>) and a root portion (<NUM>), the fiber-reinforced composite airfoil core (<NUM>) being subject to core thermal growth;
fiber-reinforced composite wrapping (<NUM>) around the root portion (<NUM>), the fiber-reinforced composite wrapping being subject to wrapping thermal growth; and
characterized by a buffer layer (<NUM>) between the root portion (<NUM>) and the platform (<NUM>), the buffer layer (<NUM>) absorbing a mismatch between the core thermal growth and the platform thermal growth.