US20190137101A1 - Combustor assembly for a turbine engine - Google Patents
Combustor assembly for a turbine engine Download PDFInfo
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- US20190137101A1 US20190137101A1 US16/231,103 US201816231103A US2019137101A1 US 20190137101 A1 US20190137101 A1 US 20190137101A1 US 201816231103 A US201816231103 A US 201816231103A US 2019137101 A1 US2019137101 A1 US 2019137101A1
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- Prior art keywords
- liner
- cap
- forward end
- annular dome
- combustor assembly
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/60—Support structures; Attaching or mounting means
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00012—Details of sealing devices
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00017—Assembling combustion chamber liners or subparts
Definitions
- the present subject matter relates generally to a gas turbine engine, or more particularly to a combustor assembly for a gas turbine engine.
- 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 general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section.
- 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 gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
- non-traditional high temperature materials such as ceramic matrix composite (CMC) materials
- CMC ceramic matrix composite
- an inner liner and an outer liner of gas turbine engines are more commonly being formed of CMC materials.
- CMC materials have different coefficients of thermal expansion than the traditional metal materials. Accordingly, coupling the CMC materials to the traditional metal materials can be problematic. For example, special care must be taken in attaching the inner liner and outer liner to a metallic inner dome structure and a metallic outer dome structure, respectively.
- certain gas turbine engines having the inner and outer liners formed of CMC materials have difficulty in controlling an amount of high-pressure air that flows through one or more connection points—e.g., between the inner liner and inner dome structure and the outer liner and outer dome structure—into a combustion chamber at least partially defined by the inner and outer liners.
- a combustor assembly having one more features allowing for a CMC liner to be attached to a respective metallic dome structure at an attachment point while controlling an amount of airflow therethrough would be useful. More particularly, a combustor assembly having one more features allowing for a CMC liner to be attached to a respective metallic dome structure at an attachment point while controlling an amount of airflow therethrough and allowing for relative thermal expansion would be particularly beneficial.
- a combustor assembly for a gas turbine engine.
- the combustor assembly defines an axial direction and includes a liner at least partially defining a combustion chamber.
- the liner extends between an aft end and a forward end generally along the axial direction.
- the combustor assembly also includes an annular dome including an enclosed surface defining a slot for receipt of the forward end of the liner.
- the combustor assembly also includes a cap positioned at the forward end of the liner and at least partially positioned within the slot defined by the enclosed surface of the annular dome.
- the cap includes a surface configured to contact at least one of the enclosed surface of the annular dome or the forward end of the liner.
- a cap assembly for a liner of a gas turbine engine combustor assembly includes a first arm and a second arm extending substantially parallel with the first arm. The first and second arms together define an opening for receipt of a forward end of the liner.
- the cap assembly also includes a base extending between the first and second arms and defining an inside surface and an outside surface.
- the cap assembly also includes a resilient member positioned adjacent to the inside surface of the base for pressing the base away from the forward end of the liner and forming a seal between the base and the forward end of the liner when the cap assembly is positioned over the forward end of the liner.
- a gas turbine engine defining an axial direction.
- the gas turbine engine includes a compressor section, a turbine section mechanically coupled to the compressor section through a shaft, and a combustor assembly disposed between the compressor section and the turbine section.
- the combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end generally along the axial direction.
- the combustor assembly also includes an annular dome including an enclosed surface defining a slot for receipt of the forward end of the liner.
- the combustor assembly also includes a cap positioned at the forward end of the liner and at least partially positioned within the slot defined by the enclosed surface of the annular dome.
- the cap includes a surface configured to contact at least one of the enclosed surface of the annular dome or the forward end of the liner.
- FIG. 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter.
- FIG. 2 is a perspective, cross-sectional view of a combustor assembly in accordance with an exemplary embodiment of the present disclosure.
- FIG. 3 is a schematic, cross-sectional view of the exemplary combustor assembly of FIG. 2 .
- FIG. 4 is a close up, cross-sectional view of an attachment point of the exemplary combustor assembly of FIG. 2 , where a forward end of an outer liner is attached to an outer annular dome.
- FIG. 5 is a close-up, cross-sectional view of an attachment point of a combustor assembly in accordance with another exemplary embodiment of the present disclosure, where a forward end of an outer liner is attached to an outer annular dome.
- upstream and downstream refer to the relative direction with respect to fluid flow in a fluid pathway.
- upstream refers to the direction from which the fluid flows
- downstream refers to the direction to which the fluid flows.
- FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine 10 , referred to herein as “turbofan engine 10 .” As shown in FIG. 1 , the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. In general, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14 .
- 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 .
- the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner.
- the fan blades 40 extend outwardly from disk 42 generally along the radial direction R.
- Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operatively coupled to a suitable actuation member 44 configured to collectively vary the pitch of the fan blades 40 in unison.
- the fan blades 40 , disk 42 , and actuation member 44 are together rotatable about the longitudinal axis 12 by LP shaft 36 across a power gear box 46 .
- the power gear box 46 includes a plurality of gears for stepping down the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.
- the disk 42 is covered by rotatable front nacelle 48 aerodynamically contoured to promote an airflow through the plurality of fan blades 40 .
- the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the fan 38 and/or at least a portion of the core turbine engine 16 .
- the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially-spaced outlet guide vanes 52 .
- a downstream section 54 of the nacelle 50 may extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.
- a volume of air 58 enters the turbofan 10 through an associated inlet 60 of the nacelle 50 and/or fan section 14 .
- 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 arrow 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 .
- HP high pressure
- 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 .
- FIG. 2 provides a perspective, cross-sectional view of a combustor assembly 100 , which may be positioned in the combustion section 26 of the exemplary turbofan engine 10 of FIG. 1 , in accordance with an exemplary embodiment of the present disclosure
- FIG. 3 provides a side, cross-sectional view of the exemplary combustor assembly 100 of FIG. 2 .
- FIG. 2 provides a perspective, cross-sectional view of the combustor assembly 100 having an outer combustor casing 136 removed for clarity.
- the combustor assembly 100 generally includes an inner liner 102 extending between and aft end 104 and a forward end 106 generally along the axial direction A, as well as an outer liner 108 also extending between and aft end 110 and a forward end 112 generally along the axial direction A.
- the inner and outer liners 102 , 108 together at least partially define a combustion chamber 114 therebetween.
- the inner and outer liners 102 , 108 are each attached to an annular dome. More particularly, the combustor assembly 100 includes an inner annular dome 116 attached to the forward end 106 of the inner liner 102 and an outer annular dome 118 attached to the forward end 112 of the outer liner 108 .
- the inner and outer annular domes 116 , 118 each include an enclosed surface 120 defining a slot 122 for receipt of the forward ends 106 , 112 of the respective inner and outer liners 102 , 108 .
- the combustor assembly 100 further includes a plurality of fuel air mixers 126 ( FIG. 3 ) spaced along a circumferential direction C within the outer dome 118 . More particularly, the plurality of fuel air mixers 126 are disposed between the outer dome 118 and the inner dome 116 along the radial direction R. Compressed air from the compressor section of the turbofan engine 10 flows into or through the fuel air mixers 126 , where the compressed air is mixed with fuel and ignited to create the combustion gases 66 within the combustion chamber 114 .
- the inner and outer domes 116 , 118 are configured to assist in providing such a flow of compressed air from the compressor section into or through the fuel air mixers 126 .
- the outer dome 118 includes an outer cowl 126 at a forward end 128 and the inner dome 116 similarly includes an inner cowl 130 at a forward end 132 .
- the outer cowl 126 and inner cowl 130 may assist in directing the flow of compressed air from the compressor section 26 into or through one or more of the fuel air mixers 126 .
- the inner and outer domes 116 , 118 each include attachment portions configured to assist in mounting the combustor assembly 100 within the turbofan engine 10 .
- the outer dome 118 includes an attachment extension 134 configured to be mounted to an outer combustor casing 136 ( FIG. 3 ) and the inner dome 116 includes a similar attachment extension 138 configured to attach to an annular support member 140 ( FIG. 3 ) within the turbofan engine 10 .
- the inner dome 116 may be formed integrally as a single annular component, and similarly, the outer dome 118 may also be formed integrally as a single annular component.
- the inner dome 116 and/or the outer dome 118 may alternatively be formed by one or more components joined in any suitable manner.
- the outer cowl 126 may be formed separately from the outer dome 118 and attached to the forward end 128 of the outer dome 118 using, e.g., a welding process.
- the attachment extension 134 may also be formed separately from the outer dome 118 and attached to the forward end 128 of the outer dome 118 using, e.g., a welding process.
- the inner dome 116 may have a similar configuration.
- the exemplary combustor assembly 100 further includes a plurality of heat shields 142 positioned around each fuel air mixer 124 , arrange circumferentially.
- the heat shields 142 are attached to and extend between the outer dome 118 and the inner dome 116 .
- the heat shields 142 are configured to protect certain components of the turbofan engine 10 from the relatively extreme temperatures of the combustion chamber 114 .
- the inner liner 102 and outer liner 108 are each comprised of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability.
- CMC ceramic matrix composite
- Exemplary CMC materials utilized for such liners 102 , 108 may include silicon carbide, 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).
- CMC materials may have coefficients of thermal expansion in the range of about 1.3 ⁇ 10 ⁇ 6 in/in/° F. to about
- the inner dome 116 and outer dome 118 may be formed of a metal, such as a nickel-based superalloy (having a coefficient of thermal expansion of about 8.3-8.5 ⁇ 10 ⁇ 6 in/in/° F. in a temperature of approximately 1000-1200° F.) or cobalt-based superalloy (having a coefficient of thermal expansion of about 7.8-8.1 ⁇ 10 ⁇ 6 in/in/° F. in a temperature of approximately 1000-1200° F.).
- the inner and outer liners 102 , 108 may be better able to handle the extreme temperature environment presented in the combustion chamber 114 .
- a plurality of specially designed mounting assemblies 144 are utilized to attach the forward end 106 of the inner liner 102 to the inner dome 116 , as well as to attach the forward end 112 of the outer liner 108 to the outer dome 118 .
- the mounting assemblies 144 are configured to accommodate the relative thermal expansion between the inner and outer domes 116 , 118 and the inner and outer liners 102 , 108 , respectively, along the radial direction R.
- the combustor assembly 100 includes an inner piston ring 146 and an outer piston ring 148 , respectively.
- the inner piston ring 146 is attached to an inner piston ring holder 150 extending from and attached to an interior casing (which for the embodiment depicted is the annular support member 140 ).
- the outer piston ring 148 is attached to an outer piston ring holder 152 extending from and attached to an outer casing (which for the embodiment depicted includes the outer combustor casing 136 and an outer turbine casing 154 ).
- the inner piston ring holder 150 and the outer piston ring holder 152 are configured to accommodate an expansion of the inner liner 102 and the outer liner 108 generally along the axial direction A, as well as generally along the radial direction R.
- the above configuration may allow for the relative thermal expansions of the inner and outer liners 102 , 108 , formed of a CMC material, and the inner and outer domes 116 , 118 , formed of a metal material, while controlling an airflow of relatively high pressure compressed air from the compressor section 26 into the relatively low pressure combustion chamber 114 .
- such a configuration may control an airflow of relatively high pressure compressed air in a high pressure plenum 156 defined between the outer liner 108 and the outer combustor casing 136 into the relatively low pressure combustion chamber 114 , as well as an airflow of relatively high pressure compressed air in an inner passage 158 positioned radially inward from the inner liner 102 into the relatively low pressure combustion chamber 114 .
- combustion gases 66 flow from the combustion chamber 114 into and through the turbine section of the turbofan engine 10 where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of turbine stator vanes and turbine rotor blades.
- a stage 1 turbine blade 160 is depicted schematically in FIG. 3 , aft of the compressor assembly 100 .
- FIG. 4 a close up, cross-sectional view of an attachment point where the forward end 112 of the outer liner 108 is attached to the outer annular dome 118 is depicted, taken along Circle 4 - 4 of FIG. 3 .
- the mounting assemblies 144 are provided extending through the slots 122 defined by the enclosed surfaces 120 of the inner and outer annular domes 116 , 118 . More particularly, referring specifically to the outer dome 118 and forward end 112 of the outer liner 108 depicted in FIG. 4 , the outer dome 118 includes a base plate 162 and a yolk 164 . The base plate 162 and the yolk 164 each extend substantially parallel to one another, which for the embodiment depicted is a direction substantially parallel to the axial direction A of the turbofan engine 10 .
- the yolk 164 may extend circumferentially with the outer dome 118 , tracking the base plate 162 .
- the slot 122 may be considered an annular slot.
- the yolk 164 may include a plurality of circumferentially spaced tabs (see FIG. 2 ), each of the individual tabs of the yolk 164 defining individual segmented portions of the slot 122 with the base plate 162 .
- the exemplary mounting assembly 144 depicted extends through the yolk 164 of the outer dome 118 , through the forward end 112 of the outer liner 108 (positioned in the slot 122 defined by the outer dome 118 ), and through the base plate 162 of the outer dome 118 .
- the mounting assembly 144 includes a pin 166 and a bushing 168 .
- the pin 166 includes a head 170 and a body 172 , the body 172 extending through the yolk 164 , the forward end 112 of the outer liner 108 (positioned in the slot 122 ), and the base plate 162 .
- a nut 174 is attached to a distal end of the body 172 of the pin 166 .
- the pin 166 may be configured as a bolt and the nut 174 may be rotatably engaged with the pin 166 for tightening the mounting assembly 144 .
- the pen 166 and nut 174 may have any other suitable configuration.
- the pin 166 may include a body 172 defining a substantially smooth cylindrical shape and the nut 174 may be configured as a clip.
- the bushing 168 is generally cylindrical in shape and positioned around the body 172 of the pin 166 within the slot 122 .
- the bushing 168 is pressed between the yolk 164 and the base plate 162 .
- the mounting assembly 144 includes a metal grommet 176 positioned around the bushing 168 within an opening defined in the forward end 112 of the outer liner 108 .
- the metal grommet 176 may reduce an amount of wear on the forward end 112 of the outer liner 108 as the outer liner 108 moves inwardly and outwardly generally along the radial direction R relative to the outer dome 118 . More particularly, the metal grommet 176 may reduce an amount of wear around an opening 177 in the outer liner through which the mounting assembly 144 extends.
- the exemplary combustor assembly 100 further includes a cap 178 positioned at the forward end 112 of the outer liner 108 . More particularly, for the embodiment depicted, the cap 178 is positioned over the forward end 112 of the outer liner 108 and at least partially within the slot 122 of the outer dome 118 .
- the cap 178 generally includes a first arm 180 and a second arm 182 , the second arm 182 extending substantially parallel with the first arm 180 .
- the first and second arms 180 , 182 together define an opening 184 for receipt of the forward end 112 of the outer liner 108 .
- the cap 178 also includes a base 186 extending between the first and second arms 180 , 182 and defining an inside surface 188 and an outside, or end, surface 190 .
- the first arm 180 , the second arm 182 , and the base 186 are all formed integrally for the embodiment depicted from a metal material.
- the first arm 180 , the second arm 182 , and the base 186 may be formed using a casting process, or alternatively, the opening 184 may be formed between the first and second arms 180 , 182 using an extrusion process.
- the cap 178 may be formed of individual arm and base components joined using, e.g., a welding process.
- the cap 178 may be a single annular component, or alternatively the cap 178 may be formed of a plurality of components arrange circumferentially over the forward end 112 of the outer liner 108 . Further, in still other embodiments, the cap 178 may be formed partially or completely of a suitable CMC material.
- the first and second arms 180 , 182 of the cap 178 extend past the mounting assemblies 144 .
- the first arm 180 and the second arm 182 may each define one or more openings for receiving at least a portion of one or more of the mounting assemblies 144 mounting the forward end 112 of the outer liner 108 to the outer dome 118 .
- the first and second arms 180 , 182 depicted may each define one or more openings allowing the metal grommet 176 , the bushing 168 , and the pin 166 of each mounting assembly 144 to extend therethrough.
- the base 186 of the cap 178 and the forward end 112 of the outer liner 108 define a gap 192 therebetween with a resilient member 194 positioned therein (i.e., adjacent to the inside surface 188 of the base 186 and the forward end 112 of the outer liner 108 ).
- the purpose of the resilient member 194 is twofold. First, the resilient member 194 is configured to form a seal between the inside surface 188 of the base 186 of the cap 178 and the forward end 112 of the outer liner 108 .
- the resilient member 194 is configured to press the base 186 of the cap 178 away from the forward end 112 of the liner 108 such that the end surface 190 of the cap 178 is pressed against the enclosed surface 120 of the outer dome 118 . Accordingly, such a configuration may allow the cap 178 to form a substantially airtight seal between the forward end 112 of the outer liner 108 and the outer dome 118 .
- the resilient member 194 may be a rope seal, such as a braided rope seal having a silicone core.
- any other suitable resilient member 194 may be provided for pressing the base 186 of the cap 178 away from the forward end of the liner and forming a seal between the inside surface 188 of the base 186 of the cap 178 and the forward end 112 of the liner 108 .
- the resilient member 194 may be a W-seal, a wire seal, or any other suitable seal.
- the forward end 106 of the inner liner 102 may be attached to the inner dome 116 in substantially the same manner that the forward end 112 of the outer liner 108 is attached to the outer dome 118 .
- a cap (similar to the cap 178 positioned over the forward end 112 of the outer liner 108 ) may be positioned over the forward end 106 of the inner liner 102 and at least partially within the slot 122 of the inner dome 116 .
- An end surface of such cap may contact the enclosed surface 120 of the inner dome 116 such that a substantially airtight seal is formed between the end surface of such cap and the enclosed surface 120 of the inner dome 116 .
- such a cap may define a gap (similar to the gap 192 defined between the inner surface 188 of the cap 178 and the forward end 112 of the outer liner 108 ) between an inner surface and the forward end 106 of the inner liner 102 with a suitable resilient member 194 positioned therein.
- a combustor in accordance with an exemplary embodiment of the present disclosure assembly having a cap positioned over an inner liner or an outer liner may be capable of controlling an airflow from a relatively high pressure plenum or a relatively high pressure inner passage into a combustion chamber through an attachment point between the inner or outer liners and an inner or outer dome.
- such a combustor assembly may be capable of controlling an airflow from a relatively high pressure plenum or a relatively high pressure inner passage into a combustion chamber through an attachment point between the inner or outer liners and an inner or outer dome while still accommodating a relative thermal expansion between the inner or outer liners and inner or outer domes.
- FIG. 5 provides a close-up, cross-sectional view of a combustor assembly 100 in accordance with another exemplary embodiment of the present disclosure. More particularly, FIG. 5 provides a close-up, cross-sectional view of an attachment point where a forward end 112 of an outer liner 108 is attached to an outer annular dome 118 .
- the exemplary combustor assembly 100 of FIG. 5 may be configured in substantially the same manner as the exemplary combustor assembly 100 described above with reference to FIGS. 2 through 4 . Accordingly, the same or similar numbering refers to the same or similar components.
- the forward end 112 of the outer liner 108 is positioned within a slot 122 defined by an enclosed surface 120 of the outer annular dome 118 .
- a mounting assembly 144 attaches the forward end 112 of the outer liner 108 to the outer annular dome 118 .
- the exemplary combustor assembly 100 depicted in FIG. 5 includes a cap 178 ′ positioned at the forward end 112 of the outer liner 108 and at least partially within the slot 122 defined by the enclosed surface 120 of the annular dome 118 .
- the exemplary cap 178 ′ depicted defines an inside surface 188 ′ configured to contact the forward end 112 of the liner 108 and an end surface 190 ′ positioned opposite the inside surface 188 ′.
- the end surface 190 ′ defines a notch 196 .
- a resilient member 194 is positioned adjacent to the end surface 190 ′ of the cap 178 ′, within the notch 196 .
- the cap 178 ′ and resilient member 194 are configured to form a seal between the end surface 190 ′ of the cap 178 ′ and the enclosed surface 120 of the annular dome 118 , as well as between the inside surface 188 ′ of the cap 178 ′ and the forward end 112 of the liner 108 .
- the resilient member 194 is configured to form a seal between the enclosed surface 120 of the annular dome 118 and the end surface 190 ′ of the cap 178 ′, and is also configured to press the cap 178 ′ against the forward end 112 of the outer liner 108 , such that the inside surface 188 ′ of the cap 178 ′ contacts the forward end 112 of the outer liner 108 .
- Such a configuration may allow for the exemplary combustor assembly 100 depicted to control an airflow through the attachment point between the outer annular dome 118 and the outer liner 108 as the outer annular dome 118 thermally expands relative to the outer liner 108 along the radial direction R.
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Abstract
Description
- This application is a division of U.S. application Ser. No. 14/842,883, filed on Sep. 2, 2015, titled “COMBUSTOR ASSEMBLY FOR A TURBINE ENGINE”, which is hereby expressly incorporated herein by reference in its entirety.
- The present subject matter relates generally to a gas turbine engine, or more particularly to a combustor assembly for a gas turbine engine.
- 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 general 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 gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
- More commonly, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are being used as structural components within gas turbine engines. For example, given an ability for CMC materials to withstand relatively extreme temperatures, there is particular interest in replacing components within the combustion section of the gas turbine engine with CMC materials. More particularly, an inner liner and an outer liner of gas turbine engines are more commonly being formed of CMC materials.
- However, certain gas turbine engines have had problems accommodating certain mechanical properties of the CMC materials incorporated therein. For example, CMC materials have different coefficients of thermal expansion than the traditional metal materials. Accordingly, coupling the CMC materials to the traditional metal materials can be problematic. For example, special care must be taken in attaching the inner liner and outer liner to a metallic inner dome structure and a metallic outer dome structure, respectively.
- Moreover, certain gas turbine engines having the inner and outer liners formed of CMC materials have difficulty in controlling an amount of high-pressure air that flows through one or more connection points—e.g., between the inner liner and inner dome structure and the outer liner and outer dome structure—into a combustion chamber at least partially defined by the inner and outer liners.
- Accordingly, a combustor assembly having one more features allowing for a CMC liner to be attached to a respective metallic dome structure at an attachment point while controlling an amount of airflow therethrough would be useful. More particularly, a combustor assembly having one more features allowing for a CMC liner to be attached to a respective metallic dome structure at an attachment point while controlling an amount of airflow therethrough and allowing for relative thermal expansion would be particularly beneficial.
- 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 disclosure, a combustor assembly for a gas turbine engine is provided. The combustor assembly defines an axial direction and includes a liner at least partially defining a combustion chamber. The liner extends between an aft end and a forward end generally along the axial direction. The combustor assembly also includes an annular dome including an enclosed surface defining a slot for receipt of the forward end of the liner. The combustor assembly also includes a cap positioned at the forward end of the liner and at least partially positioned within the slot defined by the enclosed surface of the annular dome. The cap includes a surface configured to contact at least one of the enclosed surface of the annular dome or the forward end of the liner.
- In another exemplary embodiment of the present disclosure, a cap assembly for a liner of a gas turbine engine combustor assembly is provided. The cap assembly includes a first arm and a second arm extending substantially parallel with the first arm. The first and second arms together define an opening for receipt of a forward end of the liner. The cap assembly also includes a base extending between the first and second arms and defining an inside surface and an outside surface. The cap assembly also includes a resilient member positioned adjacent to the inside surface of the base for pressing the base away from the forward end of the liner and forming a seal between the base and the forward end of the liner when the cap assembly is positioned over the forward end of the liner.
- In still another exemplary embodiment of the present disclosure, a gas turbine engine defining an axial direction is provided. The gas turbine engine includes a compressor section, a turbine section mechanically coupled to the compressor section through a shaft, and a combustor assembly disposed between the compressor section and the turbine section. The combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end generally along the axial direction. The combustor assembly also includes an annular dome including an enclosed surface defining a slot for receipt of the forward end of the liner. The combustor assembly also includes a cap positioned at the forward end of the liner and at least partially positioned within the slot defined by the enclosed surface of the annular dome. The cap includes a surface configured to contact at least one of the enclosed surface of the annular dome or the forward end of the liner.
- 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:
-
FIG. 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter. -
FIG. 2 is a perspective, cross-sectional view of a combustor assembly in accordance with an exemplary embodiment of the present disclosure. -
FIG. 3 is a schematic, cross-sectional view of the exemplary combustor assembly ofFIG. 2 . -
FIG. 4 is a close up, cross-sectional view of an attachment point of the exemplary combustor assembly ofFIG. 2 , where a forward end of an outer liner is attached to an outer annular dome. -
FIG. 5 is a close-up, cross-sectional view of an attachment point of a combustor assembly in accordance with another exemplary embodiment of the present disclosure, where a forward end of an outer liner is attached to an outer annular dome. - 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,
FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” As shown inFIG. 1 , theturbofan engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference) and a radial direction R. In general, theturbofan 10 includes afan section 14 and acore turbine engine 16 disposed downstream from thefan section 14. - The exemplary
core turbine engine 16 depicted generally includes a substantially tubularouter casing 18 that defines anannular inlet 20. Theouter 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; acombustion section 26; a turbine section including a high pressure (HP)turbine 28 and a low pressure (LP)turbine 30; and a jetexhaust nozzle section 32. A high pressure (HP) shaft orspool 34 drivingly connects the HPturbine 28 to the HPcompressor 24. A low pressure (LP) shaft orspool 36 drivingly connects theLP turbine 30 to theLP compressor 22. - For the embodiment depicted, the
fan section 14 includes avariable pitch fan 38 having a plurality offan blades 40 coupled to adisk 42 in a spaced apart manner. As depicted, thefan blades 40 extend outwardly fromdisk 42 generally along the radial direction R. Eachfan blade 40 is rotatable relative to thedisk 42 about a pitch axis P by virtue of thefan blades 40 being operatively coupled to asuitable actuation member 44 configured to collectively vary the pitch of thefan blades 40 in unison. Thefan blades 40,disk 42, andactuation member 44 are together rotatable about thelongitudinal axis 12 byLP shaft 36 across apower gear box 46. Thepower gear box 46 includes a plurality of gears for stepping down the rotational speed of theLP shaft 36 to a more efficient rotational fan speed. - Referring still to the exemplary embodiment of
FIG. 1 , thedisk 42 is covered byrotatable front nacelle 48 aerodynamically contoured to promote an airflow through the plurality offan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing orouter nacelle 50 that circumferentially surrounds thefan 38 and/or at least a portion of thecore turbine engine 16. It should be appreciated that thenacelle 50 may be configured to be supported relative to thecore turbine engine 16 by a plurality of circumferentially-spaced outlet guide vanes 52. Moreover, adownstream section 54 of thenacelle 50 may extend over an outer portion of thecore turbine engine 16 so as to define abypass airflow passage 56 therebetween. - During operation of the
turbofan engine 10, a volume ofair 58 enters theturbofan 10 through an associatedinlet 60 of thenacelle 50 and/orfan section 14. As the volume ofair 58 passes across thefan blades 40, a first portion of theair 58 as indicated byarrows 62 is directed or routed into thebypass airflow passage 56 and a second portion of theair 58 as indicated byarrow 64 is directed or routed into theLP compressor 22. The ratio between the first portion ofair 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the second portion ofair 64 is then increased as it is routed through the high pressure (HP)compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66. - The
combustion gases 66 are routed through theHP turbine 28 where a portion of thermal and/or kinetic energy from thecombustion gases 66 is extracted via sequential stages of HPturbine stator vanes 68 that are coupled to theouter casing 18 and HPturbine rotor blades 70 that are coupled to the HP shaft orspool 34, thus causing the HP shaft orspool 34 to rotate, thereby supporting operation of theHP compressor 24. Thecombustion gases 66 are then routed through theLP turbine 30 where a second portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LPturbine stator vanes 72 that are coupled to theouter casing 18 and LPturbine rotor blades 74 that are coupled to the LP shaft orspool 36, thus causing the LP shaft orspool 36 to rotate, thereby supporting operation of theLP compressor 22 and/or rotation of thefan 38. - The
combustion gases 66 are subsequently routed through the jetexhaust nozzle section 32 of thecore turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion ofair 62 is substantially increased as the first portion ofair 62 is routed through thebypass airflow passage 56 before it is exhausted from a fannozzle exhaust section 76 of theturbofan 10, also providing propulsive thrust. TheHP turbine 28, theLP turbine 30, and the jetexhaust nozzle section 32 at least partially define ahot gas path 78 for routing thecombustion gases 66 through thecore turbine engine 16. - Referring now to
FIGS. 2 and 3 , close-up cross-sectional views are provided of thecombustion section 26 of theexemplary turbofan engine 10 ofFIG. 1 . More particularly,FIG. 2 provides a perspective, cross-sectional view of acombustor assembly 100, which may be positioned in thecombustion section 26 of theexemplary turbofan engine 10 ofFIG. 1 , in accordance with an exemplary embodiment of the present disclosure, andFIG. 3 provides a side, cross-sectional view of theexemplary combustor assembly 100 ofFIG. 2 . Notably,FIG. 2 provides a perspective, cross-sectional view of thecombustor assembly 100 having anouter combustor casing 136 removed for clarity. - As shown, the
combustor assembly 100 generally includes aninner liner 102 extending between andaft end 104 and aforward end 106 generally along the axial direction A, as well as anouter liner 108 also extending between andaft end 110 and aforward end 112 generally along the axial direction A. The inner andouter liners combustion chamber 114 therebetween. The inner andouter liners combustor assembly 100 includes an innerannular dome 116 attached to theforward end 106 of theinner liner 102 and an outerannular dome 118 attached to theforward end 112 of theouter liner 108. As will be discussed in greater detail below, the inner and outerannular domes enclosed surface 120 defining aslot 122 for receipt of the forward ends 106, 112 of the respective inner andouter liners - The
combustor assembly 100 further includes a plurality of fuel air mixers 126 (FIG. 3 ) spaced along a circumferential direction C within theouter dome 118. More particularly, the plurality offuel air mixers 126 are disposed between theouter dome 118 and theinner dome 116 along the radial direction R. Compressed air from the compressor section of theturbofan engine 10 flows into or through thefuel air mixers 126, where the compressed air is mixed with fuel and ignited to create thecombustion gases 66 within thecombustion chamber 114. The inner andouter domes fuel air mixers 126. For example, theouter dome 118 includes anouter cowl 126 at aforward end 128 and theinner dome 116 similarly includes aninner cowl 130 at aforward end 132. Theouter cowl 126 andinner cowl 130 may assist in directing the flow of compressed air from thecompressor section 26 into or through one or more of thefuel air mixers 126. - Moreover, the inner and
outer domes combustor assembly 100 within theturbofan engine 10. For example, theouter dome 118 includes anattachment extension 134 configured to be mounted to an outer combustor casing 136 (FIG. 3 ) and theinner dome 116 includes asimilar attachment extension 138 configured to attach to an annular support member 140 (FIG. 3 ) within theturbofan engine 10. In certain exemplary embodiments, theinner dome 116 may be formed integrally as a single annular component, and similarly, theouter dome 118 may also be formed integrally as a single annular component. It should be appreciated, however, that in other exemplary embodiments, theinner dome 116 and/or theouter dome 118 may alternatively be formed by one or more components joined in any suitable manner. For example, with reference to theouter dome 118, in certain exemplary embodiments, theouter cowl 126 may be formed separately from theouter dome 118 and attached to theforward end 128 of theouter dome 118 using, e.g., a welding process. Similarly, theattachment extension 134 may also be formed separately from theouter dome 118 and attached to theforward end 128 of theouter dome 118 using, e.g., a welding process. Additionally, or alternatively, theinner dome 116 may have a similar configuration. - Referring still to
FIGS. 2 and 3 , theexemplary combustor assembly 100 further includes a plurality ofheat shields 142 positioned around eachfuel air mixer 124, arrange circumferentially. Theheat shields 142, for the embodiment depicted, are attached to and extend between theouter dome 118 and theinner dome 116. Theheat shields 142 are configured to protect certain components of theturbofan engine 10 from the relatively extreme temperatures of thecombustion chamber 114. - For the embodiment depicted, the
inner liner 102 andouter liner 108 are each comprised of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized forsuch liners - By contrast, the
inner dome 116 andouter dome 118, including theinner cowl 130 andouter cowl 126, respectively, may be formed of a metal, such as a nickel-based superalloy (having a coefficient of thermal expansion of about 8.3-8.5×10−6 in/in/° F. in a temperature of approximately 1000-1200° F.) or cobalt-based superalloy (having a coefficient of thermal expansion of about 7.8-8.1×10−6 in/in/° F. in a temperature of approximately 1000-1200° F.). Thus, the inner andouter liners combustion chamber 114. However, attaching the inner andouter liners outer domes assemblies 144 are utilized to attach theforward end 106 of theinner liner 102 to theinner dome 116, as well as to attach theforward end 112 of theouter liner 108 to theouter dome 118. The mountingassemblies 144 are configured to accommodate the relative thermal expansion between the inner andouter domes outer liners - Referring particularly to
FIG. 3 , at theaft end 104 of theinner liner 102 and at theaft end 110 of theouter liner 108, thecombustor assembly 100 includes aninner piston ring 146 and anouter piston ring 148, respectively. Theinner piston ring 146 is attached to an innerpiston ring holder 150 extending from and attached to an interior casing (which for the embodiment depicted is the annular support member 140). Similarly, theouter piston ring 148 is attached to an outerpiston ring holder 152 extending from and attached to an outer casing (which for the embodiment depicted includes theouter combustor casing 136 and an outer turbine casing 154). The innerpiston ring holder 150 and the outerpiston ring holder 152 are configured to accommodate an expansion of theinner liner 102 and theouter liner 108 generally along the axial direction A, as well as generally along the radial direction R. - As will be discussed in greater detail below, the above configuration may allow for the relative thermal expansions of the inner and
outer liners outer domes compressor section 26 into the relatively lowpressure combustion chamber 114. More particularly, such a configuration may control an airflow of relatively high pressure compressed air in ahigh pressure plenum 156 defined between theouter liner 108 and theouter combustor casing 136 into the relatively lowpressure combustion chamber 114, as well as an airflow of relatively high pressure compressed air in aninner passage 158 positioned radially inward from theinner liner 102 into the relatively lowpressure combustion chamber 114. - Referring particularly to
FIG. 3 , and as is discussed above, thecombustion gases 66 flow from thecombustion chamber 114 into and through the turbine section of theturbofan engine 10 where a portion of thermal and/or kinetic energy from thecombustion gases 66 is extracted via sequential stages of turbine stator vanes and turbine rotor blades. A stage 1turbine blade 160 is depicted schematically inFIG. 3 , aft of thecompressor assembly 100. - Referring now particularly to
FIG. 4 , a close up, cross-sectional view of an attachment point where theforward end 112 of theouter liner 108 is attached to the outerannular dome 118 is depicted, taken along Circle 4-4 ofFIG. 3 . - As stated, to allow for a relative thermal expansion of the
outer liner 108 andouter dome 118, the mountingassemblies 144 are provided extending through theslots 122 defined by theenclosed surfaces 120 of the inner and outerannular domes outer dome 118 andforward end 112 of theouter liner 108 depicted inFIG. 4 , theouter dome 118 includes abase plate 162 and ayolk 164. Thebase plate 162 and theyolk 164 each extend substantially parallel to one another, which for the embodiment depicted is a direction substantially parallel to the axial direction A of theturbofan engine 10. Additionally, in certain exemplary embodiments, theyolk 164 may extend circumferentially with theouter dome 118, tracking thebase plate 162. With such a configuration, theslot 122 may be considered an annular slot. However, in other embodiments, theyolk 164 may include a plurality of circumferentially spaced tabs (seeFIG. 2 ), each of the individual tabs of theyolk 164 defining individual segmented portions of theslot 122 with thebase plate 162. - The exemplary mounting
assembly 144 depicted extends through theyolk 164 of theouter dome 118, through theforward end 112 of the outer liner 108 (positioned in theslot 122 defined by the outer dome 118), and through thebase plate 162 of theouter dome 118. For the embodiment depicted, the mountingassembly 144 includes apin 166 and abushing 168. Thepin 166 includes ahead 170 and abody 172, thebody 172 extending through theyolk 164, theforward end 112 of the outer liner 108 (positioned in the slot 122), and thebase plate 162. Anut 174 is attached to a distal end of thebody 172 of thepin 166. In certain exemplary embodiments, thepin 166 may be configured as a bolt and thenut 174 may be rotatably engaged with thepin 166 for tightening the mountingassembly 144. Alternatively, however, in other exemplary embodiments, thepen 166 andnut 174 may have any other suitable configuration. For example, in other exemplary embodiments, thepin 166 may include abody 172 defining a substantially smooth cylindrical shape and thenut 174 may be configured as a clip. - Additionally, the
bushing 168 is generally cylindrical in shape and positioned around thebody 172 of thepin 166 within theslot 122. Thebushing 168 is pressed between theyolk 164 and thebase plate 162. Moreover, for the embodiment depicted, the mountingassembly 144 includes ametal grommet 176 positioned around thebushing 168 within an opening defined in theforward end 112 of theouter liner 108. Themetal grommet 176 may reduce an amount of wear on theforward end 112 of theouter liner 108 as theouter liner 108 moves inwardly and outwardly generally along the radial direction R relative to theouter dome 118. More particularly, themetal grommet 176 may reduce an amount of wear around an opening 177 in the outer liner through which the mountingassembly 144 extends. - Referring still to
FIG. 4 , theexemplary combustor assembly 100 further includes acap 178 positioned at theforward end 112 of theouter liner 108. More particularly, for the embodiment depicted, thecap 178 is positioned over theforward end 112 of theouter liner 108 and at least partially within theslot 122 of theouter dome 118. Thecap 178 generally includes afirst arm 180 and asecond arm 182, thesecond arm 182 extending substantially parallel with thefirst arm 180. The first andsecond arms opening 184 for receipt of theforward end 112 of theouter liner 108. Thecap 178 also includes a base 186 extending between the first andsecond arms inside surface 188 and an outside, or end,surface 190. Thefirst arm 180, thesecond arm 182, and the base 186 are all formed integrally for the embodiment depicted from a metal material. For example, thefirst arm 180, thesecond arm 182, and the base 186 may be formed using a casting process, or alternatively, theopening 184 may be formed between the first andsecond arms cap 178 may be formed of individual arm and base components joined using, e.g., a welding process. Moreover, in certain exemplary embodiments, thecap 178 may be a single annular component, or alternatively thecap 178 may be formed of a plurality of components arrange circumferentially over theforward end 112 of theouter liner 108. Further, in still other embodiments, thecap 178 may be formed partially or completely of a suitable CMC material. - Referring still to the embodiment depicted, the first and
second arms cap 178 extend past the mountingassemblies 144. Accordingly, thefirst arm 180 and thesecond arm 182 may each define one or more openings for receiving at least a portion of one or more of the mountingassemblies 144 mounting theforward end 112 of theouter liner 108 to theouter dome 118. For example, the first andsecond arms metal grommet 176, thebushing 168, and thepin 166 of each mountingassembly 144 to extend therethrough. - For the exemplary embodiment depicted, the
base 186 of thecap 178 and theforward end 112 of theouter liner 108 define agap 192 therebetween with aresilient member 194 positioned therein (i.e., adjacent to theinside surface 188 of thebase 186 and theforward end 112 of the outer liner 108). The purpose of theresilient member 194 is twofold. First, theresilient member 194 is configured to form a seal between theinside surface 188 of thebase 186 of thecap 178 and theforward end 112 of theouter liner 108. Second, theresilient member 194 is configured to press thebase 186 of thecap 178 away from theforward end 112 of theliner 108 such that theend surface 190 of thecap 178 is pressed against theenclosed surface 120 of theouter dome 118. Accordingly, such a configuration may allow thecap 178 to form a substantially airtight seal between theforward end 112 of theouter liner 108 and theouter dome 118. - In certain exemplary embodiments, the
resilient member 194 may be a rope seal, such as a braided rope seal having a silicone core. Alternatively, however, in other exemplary embodiments, any other suitableresilient member 194 may be provided for pressing thebase 186 of thecap 178 away from the forward end of the liner and forming a seal between theinside surface 188 of thebase 186 of thecap 178 and theforward end 112 of theliner 108. For example, in other exemplary embodiments, theresilient member 194 may be a W-seal, a wire seal, or any other suitable seal. - Moreover, referring back to
FIG. 3 , it should be appreciated that theforward end 106 of theinner liner 102 may be attached to theinner dome 116 in substantially the same manner that theforward end 112 of theouter liner 108 is attached to theouter dome 118. For example, a cap (similar to thecap 178 positioned over theforward end 112 of the outer liner 108) may be positioned over theforward end 106 of theinner liner 102 and at least partially within theslot 122 of theinner dome 116. An end surface of such cap may contact theenclosed surface 120 of theinner dome 116 such that a substantially airtight seal is formed between the end surface of such cap and theenclosed surface 120 of theinner dome 116. Further, such a cap may define a gap (similar to thegap 192 defined between theinner surface 188 of thecap 178 and theforward end 112 of the outer liner 108) between an inner surface and theforward end 106 of theinner liner 102 with a suitableresilient member 194 positioned therein. - A combustor in accordance with an exemplary embodiment of the present disclosure assembly having a cap positioned over an inner liner or an outer liner may be capable of controlling an airflow from a relatively high pressure plenum or a relatively high pressure inner passage into a combustion chamber through an attachment point between the inner or outer liners and an inner or outer dome. Moreover, such a combustor assembly may be capable of controlling an airflow from a relatively high pressure plenum or a relatively high pressure inner passage into a combustion chamber through an attachment point between the inner or outer liners and an inner or outer dome while still accommodating a relative thermal expansion between the inner or outer liners and inner or outer domes.
- Reference will now be made to
FIG. 5 .FIG. 5 provides a close-up, cross-sectional view of acombustor assembly 100 in accordance with another exemplary embodiment of the present disclosure. More particularly,FIG. 5 provides a close-up, cross-sectional view of an attachment point where aforward end 112 of anouter liner 108 is attached to an outerannular dome 118. Theexemplary combustor assembly 100 ofFIG. 5 may be configured in substantially the same manner as theexemplary combustor assembly 100 described above with reference toFIGS. 2 through 4 . Accordingly, the same or similar numbering refers to the same or similar components. - As is depicted, the
forward end 112 of theouter liner 108 is positioned within aslot 122 defined by anenclosed surface 120 of the outerannular dome 118. A mountingassembly 144 attaches theforward end 112 of theouter liner 108 to the outerannular dome 118. Additionally, theexemplary combustor assembly 100 depicted inFIG. 5 includes acap 178′ positioned at theforward end 112 of theouter liner 108 and at least partially within theslot 122 defined by theenclosed surface 120 of theannular dome 118. Theexemplary cap 178′ depicted defines aninside surface 188′ configured to contact theforward end 112 of theliner 108 and anend surface 190′ positioned opposite theinside surface 188′. Theend surface 190′ defines anotch 196. Aresilient member 194 is positioned adjacent to theend surface 190′ of thecap 178′, within thenotch 196. Thecap 178′ andresilient member 194 are configured to form a seal between theend surface 190′ of thecap 178′ and theenclosed surface 120 of theannular dome 118, as well as between theinside surface 188′ of thecap 178′ and theforward end 112 of theliner 108. More particularly, theresilient member 194 is configured to form a seal between theenclosed surface 120 of theannular dome 118 and theend surface 190′ of thecap 178′, and is also configured to press thecap 178′ against theforward end 112 of theouter liner 108, such that theinside surface 188′ of thecap 178′ contacts theforward end 112 of theouter liner 108. Such a configuration may allow for theexemplary combustor assembly 100 depicted to control an airflow through the attachment point between the outerannular dome 118 and theouter liner 108 as the outerannular dome 118 thermally expands relative to theouter liner 108 along the radial direction R. - 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 languages of the claims.
Claims (20)
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US16/231,103 US20190137101A1 (en) | 2015-09-02 | 2018-12-21 | Combustor assembly for a turbine engine |
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US16/231,103 US20190137101A1 (en) | 2015-09-02 | 2018-12-21 | Combustor assembly for a turbine engine |
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-
2015
- 2015-09-02 US US14/842,883 patent/US10197278B2/en active Active
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2016
- 2016-08-23 JP JP2016162308A patent/JP2017049000A/en active Pending
- 2016-08-25 CA CA2940025A patent/CA2940025A1/en not_active Abandoned
- 2016-08-26 EP EP16185947.5A patent/EP3139089A1/en not_active Withdrawn
- 2016-08-31 CN CN201610776933.9A patent/CN106482156B/en active Active
- 2016-08-31 CN CN201811532207.8A patent/CN110043923B/en active Active
-
2018
- 2018-12-21 US US16/231,103 patent/US20190137101A1/en not_active Abandoned
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US11859819B2 (en) | 2021-10-15 | 2024-01-02 | General Electric Company | Ceramic composite combustor dome and liners |
Also Published As
Publication number | Publication date |
---|---|
CN106482156B (en) | 2020-04-28 |
CN106482156A (en) | 2017-03-08 |
JP2017049000A (en) | 2017-03-09 |
CN110043923A (en) | 2019-07-23 |
US20170059160A1 (en) | 2017-03-02 |
CN110043923B (en) | 2021-08-20 |
CA2940025A1 (en) | 2017-03-02 |
EP3139089A1 (en) | 2017-03-08 |
US10197278B2 (en) | 2019-02-05 |
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