Patent Grant
10663167
The present subject matter relates generally to combustor assemblies for gas turbine engines. More particularly, the present subject matter relates to combustor assemblies utilizing ceramic matrix composite combustor domes.
A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
Combustion gas temperatures are relatively hot, such that some components in or near the combustion section and the downstream turbine section require features for deflecting or mitigating the effects of the combustion gas temperatures. More commonly, non-traditional high temperature composite materials, such as ceramic matrix composite (CMC) materials, are being used in applications such as gas turbine engine combustion and turbine sections. Components fabricated from CMC materials have a higher temperature capability compared with typical components, e.g., metal components, which may allow improved component performance and/or increased system temperatures. Often, components in direct contact with the hot combustion gases may be fabricated from a CMC material, while combustor assembly support structures comprise metallic components, which are less capable of withstanding high temperatures than CMC components and have different coefficients of thermal expansion (CTE) than CMC components. Therefore, exposing the metallic support structure to the relatively high combustion temperatures risks overheating the metallic support structure and the CTE mismatch between the metallic and CMC components can place undue thermal stresses on CMC components mounted to the metallic support structure.
Accordingly, improved combustion assemblies for mitigating the negative effects of using CMC components with metallic hardware would be desirable. As an example, a combustor assembly having a CMC combustor dome that shields a metallic support structure from a combustion chamber of the combustor assembly would be beneficial. As another example, a combustor assembly that decouples a CMC combustor dome from a structural load path of the combustor assembly would be advantageous. Additionally, a CMC combustor dome formed from a plurality of CMC tiles, e.g., to simplify manufacturing and repair of the dome while also reducing unacceptable natural frequencies of the dome, would be desirable.
Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.
In one exemplary embodiment of the present subject matter, a combustor assembly is provided. The combustor assembly comprises an annular ceramic matrix composite (CMC) inner liner including an inner liner flange, an annular CMC outer liner radially spaced apart from the CMC inner liner and including an outer liner flange, and an annular CMC combustor dome comprising a plurality of tiles positioned circumferentially adjacent one another. Each tile of the plurality of tiles has a first end radially opposite a second end. The CMC inner liner, the CMC outer liner, and the CMC combustor dome form a combustor that defines a combustion chamber. The CMC combustor dome is positioned at a forward end of the combustor. The combustor assembly also comprises a support structure for supporting the combustor. The support structure includes an annular frame having a frame channel defining a groove, and the first end of each tile of the plurality of tiles disposed within the groove of the frame channel. The support structure further includes an inner support flange and an outer support flange. The inner liner flange is secured to the inner support flange and the outer liner flange is secured to the outer support flange.
In another exemplary embodiment of the present subject matter, a combustor assembly is provided. The combustor assembly comprises an annular ceramic matrix composite (CMC) inner liner including an inner liner flange, an annular CMC outer liner radially spaced apart from the CMC inner liner and including an outer liner flange; and an annular CMC combustor dome having a first end and a radially opposite second end. The CMC inner liner, the CMC outer liner, and the CMC combustor dome form a combustor that defines a combustion chamber, and the CMC combustor dome is positioned at a forward end of the combustor. The combustor assembly further comprises a support structure for supporting the combustor. The support structure includes an annular frame, an inner support flange, and an outer support flange. The combustor assembly also comprises an inner CMC bracket including an inner bracket channel and an outer CMC bracket including an outer bracket channel. The first end of the CMC combustor dome is disposed within the inner bracket channel, and the second end of the CMC combustor dome is disposed within the outer bracket channel.
In a further exemplary embodiment of the present subject matter, a combustor assembly is provided. The combustor assembly comprises an annular ceramic matrix composite (CMC) inner liner including an inner liner flange, an annular CMC outer liner radially spaced apart from the CMC inner liner and including an outer liner flange, and an annular CMC combustor dome having a first end and a second end. The CMC inner liner, the CMC outer liner, and the CMC combustor dome form a combustor that defines a combustion chamber, and the CMC combustor dome is positioned at a forward end of the combustor. The combustor assembly also includes a CMC bracket including a bracket channel and a support structure for supporting the combustor. The support structure includes an annular frame including a frame channel, an inner support flange, and an outer support flange. The first end of the CMC combustor dome is disposed within the frame channel, and the second end of the CMC combustor dome is disposed within the bracket channel. Moreover, the CMC combustor dome is positioned between the support structure and the combustion chamber.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.
A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:
Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows and “downstream” refers to the direction to which the fluid flows.
Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,
The exemplary core turbine engine 16 depicted generally includes a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a combustion section 26; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22. In other embodiments of turbofan engine 10, additional spools may be provided such that engine 10 may be described as a multi-spool engine.
For the depicted embodiment, fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As depicted, fan blades 40 extend outward from disk 42 generally along the radial direction R. The fan blades 40 and disk 42 are together rotatable about the longitudinal axis 12 by LP shaft 36. In some embodiments, a power gear box having a plurality of gears may be included for stepping down the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.
Referring still to the exemplary embodiment of
During operation of the turbofan engine 10, a volume of air 58 enters turbofan 10 through an associated inlet 60 of the nacelle 50 and/or fan section 14. As the volume of air 58 passes across fan blades 40, a first portion of the air 58 as indicated by arrows 62 is directed or routed into the bypass airflow passage 56 and a second portion of the air 58 as indicated by arrows 64 is directed or routed into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly known as a bypass ratio. The pressure of the second portion of air 64 is then increased as it is routed through the high pressure (HP) compressor 24 and into the combustion section 26, where it is mixed with fuel and burned to provide combustion gases 66.
The combustion gases 66 are routed through the HP turbine 28 where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 that are coupled to the outer casing 18 and HP turbine rotor blades 70 that are coupled to the HP shaft or spool 34, thus causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 that are coupled to the outer casing 18 and LP turbine rotor blades 74 that are coupled to the LP shaft or spool 36, thus causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan 38.
The combustion gases 66 are subsequently routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is substantially increased as the first portion of air 62 is routed through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the core turbine engine 16.
It will be appreciated that, although described with respect to turbofan 10 having core turbine engine 16, the present subject matter may be applicable to other types of turbomachinery. For example, the present subject matter may be suitable for use with or in turboprops, turboshafts, turbojets, industrial and marine gas turbine engines, and/or auxiliary power units.
The inner and outer liners 102, 104 and their respective flanges 116, 118, as well as the combustor dome 114, comprise a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. As such, the inner liner 102 may be referred to as CMC inner liner 102, the outer liner 104 may be referred to as CMC outer liner 104, and the combustor dome 114 may be referred to as CMC combustor dome 114. Exemplary CMC materials and methods or techniques for forming CMC components are described in greater detail below.
The combustor 120 is supported within the gas turbine engine by a support structure 126. More particularly, the support structure 126 supports the inner and outer liners 102, 104 and the combustor dome 114, thereby supporting the combustor 120. Further, the CMC combustor dome 114 is positioned between the support structure 126 and the combustion chamber 122, such that the CMC combustor dome 114 shields the support structure 126 from direct interaction with the environment within the combustion chamber 122, such as the relatively extreme temperatures of the combustion gases 66. Accordingly, because the CMC combustor dome 114 shields the support structure 126 from the combustion chamber 122, the support structure 126 may be formed from a metallic material, such as a metal or metal alloy, which has a lower temperature capability that the CMC combustor dome 114.
As illustrated in
As shown in
Referring to
The attachment mechanisms 140 may be bolts, pins, or other suitable fasteners. Moreover, each of the inner liner flange apertures and outer liner flange apertures may include a grommet (not shown), which helps these components move radially along a bushing 142 positioned over the attachment mechanism 140 while preventing or reducing wear on the components, as well as binding of the components. The grommets may be particularly useful where the inner and outer liners 102, 104 are formed from a CMC material.
As shown in
Additionally, the combustor assembly 100 comprises an annular CMC bracket 148 that includes a bracket channel 150 defining a groove 152. The groove 152 is configured for receipt of an outer end of the combustor dome 114, as described in greater detail below. The CMC bracket 148 is secured between the outer liner flange 118 and the outer support flange 136 such that the CMC bracket 148 extends axially aft with respect to the support structure 126. It will be understood that the CMC bracket 148 defines a plurality of apertures spaced apart along the circumferential direction C, and each bracket aperture is aligned with a radial series of outer apertures such that an attachment mechanism 140 extends through each radially aligned outer support flange aperture 138d, outer liner flange aperture, bracket aperture, and outer member aperture 138c as illustrated in
In some embodiments, the CMC bracket 148 may be segmented along the circumferential direction into a plurality of CMC bracket sections that together form the annular CMC bracket 148. Thus, each bracket section includes a portion of the bracket channel 150 and defines one or more of the circumferentially spaced apart apertures for securing the bracket section with attachment mechanism(s) 140. As appropriate, one or more seals may be positioned between the circumferential edges of each bracket section, e.g., to prevent fluid leakage from the combustion chamber 122 through the crack or discontinuity formed between each bracket section.
As illustrated in
Referring still to
Further, as illustrated in
Turning now to
Each tile 164 of the plurality of tiles 164 has a first end 166 radially opposite a second end 168. As shown in
Referring back to
As further illustrated in
Turning now to
It will be understood that, for the plurality of tiles 164 forming the CMC combustor dome 114, each tile side 170, 172 may define an overlap portion 184 that overlaps with an overlap portion 184 defined by an adjacent tile side 170, 172 such that the dome 114 comprises a plurality of tiles 164 having overlapping edges. The overlapping tile edges provide a seal between each tile 164, e.g., to help prevent fluid leakage from the combustion chamber 122 through the crack or discontinuity formed between each tile 164. Of course, in other embodiments, the overlap portions 184 may be omitted such that each tile has substantially planar radial edges along sides 170, 172, and another sealing mechanism, such as a spline seal or the like, used between adjacent tile sides 170, 172 to help prevent leakage around the tiles 164.
Referring back to
In the embodiment depicted in
As described herein, the inner and outer liners 102, 104, bracket 148, and the tiles 164 forming the combustor dome 114 may be formed from a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. It may be particularly useful to utilize CMC materials in or near the hot gas path 78 due to the relatively high temperatures of the combustion gases 66, and the use of CMC materials within the combustor assembly 100 may allow reduced cooling airflow to the CMC components and higher combustion temperatures, as well as other benefits and advantages. However, other components of the turbofan engine 10, such as components of HP compressor 24, HP turbine 28, and/or LP turbine 30, also may comprise a CMC material.
Exemplary CMC materials utilized for such components may include silicon carbide (SiC), silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, bundles of the fibers, which may include a ceramic refractory material coating, are formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes may be laid up together (e.g., as plies) to form a preform component. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, such as a cure or burn-out to yield a high char residue in the preform, and subsequent chemical processing, such as melt-infiltration or chemical vapor infiltration with silicon, to arrive at a component formed of a CMC material having a desired chemical composition. In other embodiments, the CMC material may be formed as, e.g., a carbon fiber cloth rather than as a tape.
More specifically, examples of CMC materials, and particularly SiC/Si—SiC (fiber/matrix) continuous fiber-reinforced ceramic composite (CFCC) materials and processes, are described in U.S. Pat. Nos. 5,015,540; 5,330,854; 5,336,350; 5,628,938; 6,024,898; 6,258,737; 6,403,158; and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. Such processes generally entail the fabrication of CMCs using multiple pre-impregnated (prepreg) layers, e.g., the ply material may include prepreg material consisting of ceramic fibers, woven or braided ceramic fiber cloth, or stacked ceramic fiber tows that has been impregnated with matrix material. In some embodiments, each prepreg layer is in the form of a “tape” comprising the desired ceramic fiber reinforcement material, one or more precursors of the CMC matrix material, and organic resin binders. Prepreg tapes can be formed by impregnating the reinforcement material with a slurry that contains the ceramic precursor(s) and binders. Preferred materials for the precursor will depend on the particular composition desired for the ceramic matrix of the CMC component, for example, SiC powder and/or one or more carbon-containing materials if the desired matrix material is SiC. Notable carbon-containing materials include carbon black, phenolic resins, and furanic resins, including furfuryl alcohol (C4H3OCH2OH). Other typical slurry ingredients include organic binders (for example, polyvinyl butyral (PVB)) that promote the flexibility of prepreg tapes, and solvents for the binders (for example, toluene and/or methyl isobutyl ketone (MIBK)) that promote the fluidity of the slurry to enable impregnation of the fiber reinforcement material. The slurry may further contain one or more particulate fillers intended to be present in the ceramic matrix of the CMC component, for example, silicon and/or SiC powders in the case of a Si—SiC matrix. Chopped fibers or whiskers or other materials also may be embedded within the matrix as previously described. Other compositions and processes for producing composite articles, and more specifically, other slurry and prepreg tape compositions, may be used as well, such as, e.g., the processes and compositions described in U.S. Patent Application Publication No. 2013/0157037.
The resulting prepreg tape may be laid-up with other tapes, such that a CMC component formed from the tape comprises multiple laminae, each lamina derived from an individual prepreg tape. Each lamina contains a ceramic fiber reinforcement material encased in a ceramic matrix formed, wholly or in part, by conversion of a ceramic matrix precursor, e.g., during firing and densification cycles as described more fully below. In some embodiments, the reinforcement material is in the form of unidirectional arrays of tows, each tow containing continuous fibers or filaments. Alternatives to unidirectional arrays of tows may be used as well. Further, suitable fiber diameters, tow diameters, and center-to-center tow spacing will depend on the particular application, the thicknesses of the particular lamina and the tape from which it was formed, and other factors. As described above, other prepreg materials or non-prepreg materials may be used as well.
After laying up the tapes or plies to form a layup, the layup is debulked and, if appropriate, cured while subjected to elevated pressures and temperatures to produce a preform. The preform is then heated (fired) in a vacuum or inert atmosphere to decompose the binders, remove the solvents, and convert the precursor to the desired ceramic matrix material. Due to decomposition of the binders, the result is a porous CMC frame that may undergo densification, e.g., melt infiltration (MI), to fill the porosity and yield the CMC component. Specific processing techniques and parameters for the above process will depend on the particular composition of the materials. For example, silicon CMC components may be formed from fibrous material that is infiltrated with molten silicon, e.g., through a process typically referred to as the Silcomp process. Another technique of manufacturing CMC components is the method known as the slurry cast melt infiltration (MI) process. In one method of manufacturing using the slurry cast MI method, CMCs are produced by initially providing plies of balanced two-dimensional (2D) woven cloth comprising silicon carbide (SiC)-containing fibers, having two weave directions at substantially 90° angles to each other, with substantially the same number of fibers running in both directions of the weave. The term “silicon carbide-containing fiber” refers to a fiber having a composition that includes silicon carbide, and preferably is substantially silicon carbide. For instance, the fiber may have a silicon carbide core surrounded with carbon, or in the reverse, the fiber may have a carbon core surrounded by or encapsulated with silicon carbide.
Other techniques for forming CMC components include polymer infiltration and pyrolysis (PIP) and oxide/oxide processes. In PIP processes, silicon carbide fiber preforms are infiltrated with a preceramic polymer, such as polysilazane and then heat treated to form a SiC matrix. In oxide/oxide processing, aluminum or alumino-silicate fibers may be pre-impregnated and then laminated into a preselected geometry. Components may also be fabricated from a carbon fiber reinforced silicon carbide matrix (C/SiC) CMC. The C/SiC processing includes a carbon fibrous preform laid up on a tool in the preselected geometry. As utilized in the slurry cast method for SiC/SiC, the tool is made up of graphite material. The fibrous preform is supported by the tooling during a chemical vapor infiltration process at about 1200° C., whereby the C/SiC CMC component is formed. In still other embodiments, 2D, 2.5D, and/or 3D preforms may be utilized in MI, CVI, PIP, or other processes. For example, cut layers of 2D woven fabrics may be stacked in alternating weave directions as described above, or filaments may be wound or braided and combined with 3D weaving, stitching, or needling to form 2.5D or 3D preforms having multiaxial fiber architectures. Other ways of forming 2.5D or 3D preforms, e.g., using other weaving or braiding methods or utilizing 2D fabrics, may be used as well.
Thus, a variety of processes may be used to form CMC gas turbine components, such as a CMC inner liner 102, a CMC outer liner 104, a CMC bracket 148, and CMC dome tiles 164, which form a CMC combustor dome 114. Of course, other suitable processes, including variations and/or combinations of any of the processes described above, also may be used to form CMC components for use with the various retention assembly and flowpath assembly embodiments described herein.
As described herein, the present subject matter provides a combustor assembly having a CMC combustor dome that is separated from the structural load path, thereby minimizing stress and strain levels in the dome, and that shields its metallic support structure from the combustion chamber of the combustor assembly, which helps control thermal deflection and thermal stress of the metallic support structure. The combustor assembly also may utilize CMC inner and outer combustor liners. As described above, using CMC materials to form the combustor dome and liners may reduce the cooling required by the dome and liners while also allowing increased combustion temperatures, which may increase engine performance. Preferably, the CMC combustor dome is formed from a plurality of CMC dome tiles, which circumferentially segment the dome into a plurality of segments. Utilizing a plurality of CMC dome tiles rather than a single piece CMC combustor dome may simplify manufacturing of the dome as well as repair of the dome, as each dome tile may be individually replaced. Moreover, a segmented CMC combustor dome may reduce vibration within the combustor assembly by increasing the natural frequency and damping of the dome. Of course, the present subject matter may have other benefits and advantages as well.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
This invention was made with government support under contract number FA8650-15-D-2501 awarded by the U.S. Department of the Air Force. The government may have certain rights in the invention.
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