US20190203940A1 - Combustor Assembly for a Turbine Engine - Google Patents
Combustor Assembly for a Turbine Engine Download PDFInfo
- Publication number
- US20190203940A1 US20190203940A1 US15/860,835 US201815860835A US2019203940A1 US 20190203940 A1 US20190203940 A1 US 20190203940A1 US 201815860835 A US201815860835 A US 201815860835A US 2019203940 A1 US2019203940 A1 US 2019203940A1
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- Prior art keywords
- forward end
- openings
- warming
- liner
- combustor assembly
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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
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
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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
- 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
- F23R2900/00005—Preventing fatigue failures or reducing mechanical stress in gas turbine components
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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
- F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies
Definitions
- the present subject matter relates generally to gas turbine engines, and more particularly to combustor assemblies for gas turbine engines.
- 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
- inner and outer liners of gas turbine engines are more commonly being formed of CMC materials.
- a combustor assembly of a gas turbine engine that includes features that reduce the stress and strain on combustion liners of the combustor assembly during rapid power increases of the gas turbine engine would be useful.
- a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction.
- the combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end.
- the forward end of the liner defines a plurality of mounting openings spaced along the circumferential direction.
- the forward end of the liner also defines one or more blind warming openings.
- a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction.
- the combustor assembly includes a dome comprising a yolk and a base plate spaced outward from the yolk along the radial direction, the yolk and the base plate defining a slot.
- the combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot.
- the yolk defines a plurality of impingement openings extending therethrough such that a warming airflow impinges on the forward end of the liner during operation of the gas turbine engine.
- a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction.
- the combustor assembly includes a dome defining a slot and a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot of the dome, the forward end of the liner defining a plurality of mounting openings spaced along the circumferential direction.
- the combustor assembly includes a plurality of mounting assemblies for coupling the forward end of the liner with the dome, each of the plurality of mounting assemblies including: a pin extending through one of the plurality of mounting openings and a grommet positioned within one of the plurality of mounting openings and defining a grommet opening, the pin received within the grommet opening, and wherein the grommet defines one or more warming passages.
- FIG. 1 provides a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter
- FIG. 2 provides a schematic, cross-sectional view of a combustor assembly in accordance with an exemplary embodiment of the present disclosure
- FIG. 3 provides 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 dome section;
- FIG. 4 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and a plurality of blind warming openings defined by the forward end in accordance with an exemplary embodiment of the present disclosure
- FIG. 5 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and a warming slit defined by the forward end in accordance with an exemplary embodiment of the present disclosure
- FIG. 6 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and impingement openings defined by the outer dome section in accordance with an exemplary embodiment of the present disclosure
- FIG. 7 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and depicts an impingement jacket attached to the outer dome section in accordance with an exemplary embodiment of the present disclosure
- FIG. 8 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and warming passages defined by a grommet in accordance with an exemplary embodiment of the present disclosure
- FIG. 9 provides a cross sectional view of the grommet of FIG. 8 taken on line 9 - 9 of FIG. 8 ;
- FIG. 10 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and further depicting the forward end of the outer liner having enhanced surfaces in accordance with an exemplary embodiment of the present disclosure.
- 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. It should be appreciated, that as used herein, terms of approximation, such as “about” and “approximately,” refer to being within a ten percent (10%) margin of error.
- the combustor assembly includes a dome defining a slot and a liner that at least partially defines a combustion chamber.
- the liner extends between an aft end and a forward end. At least a portion of the forward end is received within the slot of the dome.
- the forward end of the liner defines a plurality of blind warming openings that allow a warming airflow to actively warm the forward end during transient operation of the engine.
- the dome defines a plurality of impingement openings that allow a warming airflow to impinge on the forward end of the liner to actively warm the forward end during transient operation of the engine.
- a grommet coupling a pin with the liner defines a plurality of warming passages that allow a warming airflow to flow therethrough to warm the forward end of the liner during transient engine operation.
- one or more surfaces of the forward end of the liner have enhanced surfaces that increase the thermal response of the forward end.
- FIG. 1 provides 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), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). 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 spinner or 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 .
- turbofan engine 10 depicted in FIG. 1 is by way of example only, and that in other exemplary embodiments, the turbofan engine 10 may have any suitable configuration.
- present disclosure matter may be suitable for use with or in turboprops, turboshafts, turbojets, reverse-flow engines, industrial and marine gas turbine engines, and/or auxiliary power units.
- FIG. 2 provides a close-up cross-sectional view of a combustor assembly 100 in accordance with an exemplary embodiment of the present disclosure.
- the combustor assembly 100 of FIG. 2 may be positioned in the combustion section 26 of the exemplary turbofan engine 10 of FIG. 1 .
- FIG. 2 provides a side, cross-sectional view of the exemplary combustor assembly 100 of FIG. 2 .
- the combustor assembly 100 includes an inner liner 102 extending between an aft end 104 and a forward end 106 generally along the axial direction A, as well as an outer liner 108 also extending between an 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 annular dome includes an inner dome section 116 attached to the forward end 106 of the inner liner 102 and an outer dome section 118 attached to the forward end 112 of the outer liner 108 .
- the inner and outer dome sections 116 , 118 may be formed integrally (or alternatively may be formed of a plurality of components attached in any suitable manner) and may each extend along the circumferential direction C to define an annular shape. As will be discussed in greater detail below with reference to FIG. 3 , the inner and outer dome sections 116 , 118 each include an inner surface 120 (i.e., inner relative to their respective forward ends) and a forward surface 121 at least partially defining a slot 122 for receipt of the forward end 106 of the inner liner 102 , and the forward end 112 of the outer liner 108 , respectively.
- an inner surface 120 i.e., inner relative to their respective forward ends
- a forward surface 121 at least partially defining a slot 122 for receipt of the forward end 106 of the inner liner 102 , and the forward end 112 of the outer liner 108 , respectively.
- the combustor assembly 100 further includes a plurality of fuel air mixers 124 spaced along a circumferential direction C and positioned at least partially within the annular dome. More particularly, the plurality of fuel air mixers 124 are disposed at least partially between the outer dome section 118 and the inner dome section 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 124 , where the compressed air is mixed with fuel and ignited to create the combustion gases 66 ( FIG. 1 ) within the combustion chamber 114 .
- the inner and outer dome sections 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 124 .
- the inner and outer dome sections 116 , 118 each include attachment portions configured to assist in mounting the combustor assembly 100 within the turbofan engine 10 ( FIG. 1 ).
- the outer dome section 118 includes an attachment extension 134 configured to be mounted to an outer combustor casing 136 and the inner dome section 116 includes a similar attachment extension 138 configured to attach to an annular support member 140 within the turbofan engine 10 .
- the inner dome section 116 may be formed integrally as a single annular component, and similarly, the outer dome section 118 may also be formed integrally as a single annular component.
- the inner dome section 116 and/or the outer dome section 118 may alternatively be formed by one or more components being joined in any suitable manner.
- the outer cowl 126 may be formed separately from the outer dome section 118 and attached to the forward end 128 of the outer dome section 118 using, e.g., a welding process.
- the attachment extension 134 may also be formed separately from the outer dome section 118 and attached to the forward end 128 of the outer dome section 118 using, e.g., a welding process.
- the inner dome section 116 may have a similar configuration.
- the exemplary combustor assembly 100 further includes a heat shield 142 positioned around the fuel air mixer 124 depicted.
- the exemplary heat shield 142 is attached to and extends between the outer dome section 118 and the inner dome section 116 .
- the heat shield 142 is configured to protect certain components of the turbofan engine 10 ( FIG. 1 ) from the relatively extreme temperatures of the combustion chamber 114 .
- the inner liner 102 and the outer liner 108 are each formed 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 annular dome including the inner dome section 116 and outer dome section 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.).
- 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.
- 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.
- combustion gases 66 flow from the combustion chamber 114 into and through the turbine section of the turbofan engine 10 ( FIG. 1 ) 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 one (1) turbine nozzle 156 is depicted schematically in FIG. 2 positioned aft of the combustor assembly 100 .
- FIG. 3 provides a close up, schematic, cross-sectional view of an attachment point where the forward end 112 of the outer liner 108 is mounted to the outer dome section 118 within the slot 122 of the outer dome section 118 .
- a plurality of mounting assemblies 144 are used to attach the outer liner 108 to the outer dome section 118 and the inner liner 102 to the inner dome section 116 . More particularly, the mounting assemblies 144 attach the forward end 112 of the outer liner 108 to the outer dome section 118 within the slot 122 of the outer dome section 118 as shown in FIGS.
- the slots 122 are defined by their respective domes. Moreover, the slots 122 receive the forward ends 106 , 112 of the inner and outer liners 102 , 108 , respectively.
- the outer dome section 118 includes a base plate 158 and a yolk 160 .
- the base plate 158 and the yolk 160 are spaced along the radial direction R.
- the base plate 158 and the yolk 160 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 (see also FIG. 2 ).
- the slot 122 is defined between the base plate 158 and the yolk 160 .
- the slot 122 is further defined by the forward surface 121 .
- the yolk 160 may extend circumferentially with the outer dome section 118 , tracking the base plate 158 .
- the slot 122 may be considered an annular slot.
- the yolk 160 may include a plurality of circumferentially spaced tabs, each of the individual tabs of the yolk 160 defining individual segmented portions of the slot 122 with the base plate 158 .
- the exemplary mounting assembly 144 depicted includes the yolk 160 of the outer dome section 118 and the base plate 158 of the outer dome section 118 . Moreover, the mounting assembly 144 includes a pin 162 and a bushing 164 .
- the pin 162 includes a head 166 and a shank 168 .
- the shank 168 extends through the yolk 160 , the forward end 112 of the outer liner 108 (positioned in slot 122 ), and the base plate 158 .
- a nut 170 is attached to a distal end of the shank 168 of the pin 162 .
- the pin 162 may be configured as a bolt and the nut 170 may be rotatably engaged with a threaded portion of the pin 162 (at, e.g., the distal end of the shank 168 ) for tightening the mounting assembly 144 .
- the pin 162 and nut 170 may have any other suitable configurations.
- the pin 162 may include a shank 168 defining a substantially smooth cylindrical shape and the nut 170 may be configured as a clip.
- the bushing 164 is generally cylindrical in shape and is positioned around the shank 168 of the pin 162 within the slot 122 .
- the bushing 164 is pressed between the yolk 160 and the base plate 158 by tightening the nut 170 on the pin 162 .
- the mounting assembly 144 includes a metal grommet 172 positioned around the bushing 164 and pin 162 .
- the grommet 172 is positioned in a mounting opening 174 defined by the forward end 112 of the outer liner 108 .
- the grommet 172 includes an outer collar 176 positioned adjacent to an outer surface 178 of the outer liner 108 and an inner collar 180 positioned adjacent to an inner surface 182 of the outer liner 108 .
- the grommet 172 additionally includes a body 184 .
- the metal grommet 172 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 section 118 .
- the mounting assembly 144 may have other suitable configurations, and further still in other embodiments, any other suitable attachment assembly may be used.
- the forward end 112 of the outer liner 108 depicted further includes an axial interface surface 186 and a radial interface surface 188 .
- the axial interface surface 186 is configured as a portion of the forward end 112 of the outer liner 108 facing the base plate 158 of the outer dome section 118 , or more particularly, facing the inner surface 120 of the outer dome section 118 .
- the radial interface surface 188 is configured as a portion of the forward end 112 of the outer liner 108 facing the forward surface 121 of the outer dome section 118 .
- the axial interface surface 186 and inner surface 120 each extend in a direction parallel to the axial direction A
- the radial interface surface 188 and forward surface 121 each extend in a direction parallel to the radial direction R.
- the axial interface surface 186 defines a radial gap G R with the inner surface 120 of the outer dome section 118 and the radial interface surface 188 defines an axial gap G A with the forward surface 121 of the outer dome section 118 .
- the combustor assembly 100 may be designed such that the radial and axial gaps G R , G A allow for only a predetermined amount of airflow therethrough into the combustion chamber 114 . Notably, allowing such a flow of air during operating conditions of the combustor assembly 100 may ensure relatively hot combustion gases within the combustion chamber 114 do not flow into and/or through the slot 122 of the outer dome section 118 , potentially damaging certain components of the combustor assembly 100 .
- additional airflow can be provided to warm the forward end 112 of the outer liner 108 (as well as the forward end 106 of the inner liner 102 depicted in FIG. 2 ) to improve the thermal response (e.g., reduce the thermal lag) of the forward ends 112 , 106 of the outer and inner liners 108 , 102 during transient operation of the turbofan engine 10 ( FIG. 1 ), and particularly during rapid power increases or bursts of the turbofan engine 10 .
- the stress and strain on the outer and inner liners 108 , 102 can be reduced during transient operation of the engine.
- FIG. 4 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118 . Further, FIG. 4 depicts the forward end 112 of the outer liner 108 defining a plurality of blind warming openings 190 .
- the diameter of the blind warming openings 190 may range between or about between 0.020 and 0.080 inches (0.508-2.032 mm).
- the diameter of the mounting openings 174 may range between or about between 0.400 and 0.800 inches (10.16-20.32 mm). Thus, the mounting openings 174 have larger diameters than the blind warming openings 190 .
- the blind warming openings 190 defined by the forward end 112 are operatively configured to increase the thermal response of the forward end 112 of the outer liner 108 . As such, as will be explained more fully below, the stress and strain on the outer liner 108 can be reduced, thereby improving the durability of the outer liner 108 .
- the forward end 106 of the inner liner 102 can likewise define blind warming openings 190 , and thus, the durability of the inner liner 102 can likewise be improved.
- each of the blind warming openings 190 extend between an opening end 192 and a terminus end 194 .
- the opening ends 192 of the blind warming openings 190 allow fluid (e.g., a warming airflow WA) to flow into and out of their respective openings and the terminus ends 194 are the dead ends or terminating points of the blind warming openings 190 .
- the blind warming openings 190 extend along the radial direction R from one of the outer surface 178 and the inner surface 182 and terminate prior to the opposing surface. In this way, the warming openings are blind openings (i.e., they do not extend through the forward end 112 of the outer liner 108 ).
- one or more of the blind warming openings 190 extend along the radial direction R from the outer surface 178 and terminate prior to the opposing inner surface 182 and one or more of the blind warming openings 190 extend along the radial direction R from the inner surface 182 and terminate prior to the opposing outer surface 178 .
- the blind warming openings 190 include outer surface openings 196 that extend along the radial direction R from the outer surface 178 and terminate prior to the inner surface 182 and the blind warming openings 190 include inner surface openings 198 that extend along the radial direction R from the inner surface 182 and terminate prior to the outer surface 178 .
- the opening ends 192 of the outer surface openings 196 are defined by the outer surface 178 of the forward end 112 and the terminus ends 194 of the outer surface openings 196 are defined by the forward end 112 at some point along the radial thickness of the forward end 112 .
- the opening ends 192 of the inner surface openings 198 are defined by the inner surface 182 of the forward end 112 and the terminus ends 194 of the inner surface openings 198 are defined by the forward end 112 at some point along the radial thickness of the forward end 112 .
- the forward end 112 defines a row of blind warming openings 190 .
- the row of blind warming openings 190 includes outer surface openings 196 that alternate with the inner surface openings 198 along the axial direction A.
- the outer surface openings 196 and the inner surface openings 198 can be axially positioned near one another without compromising the structural integrity of the forward end 112 of the outer liner 108 . In this way, efficient warming of the forward end 112 along its radial thickness and axial length can be achieved without, as noted previously, compromising the structural integrity of the forward end 112 .
- a midline is defined midway between the outer and inner surfaces 178 , 182 of the forward end 112 of the outer liner 108 along the radial direction R.
- the blind warming openings 190 extend along the radial direction R from one of the outer and inner surfaces 178 , 182 of the forward end 112 past the midline M.
- the outer surface openings 196 extend from the outer surface 178 past the midline M and the inner surface openings 198 extend from the inner surface 182 past the midline M.
- the opening ends 192 of the outer surface openings 196 extend from the outer surface 178 and their terminus ends 194 extend past the midline M.
- the blind warming openings 190 are angled with respect to the radial direction R.
- the opening ends 192 of the outer surface openings 196 and the inner surface openings 198 are positioned aft of their respective terminus ends 194 along the axial direction A.
- the opening ends 192 of the blind warming openings 190 are not aligned with their respective terminus ends 194 along the radial direction R.
- the blind warming openings 190 are angled with respect to the radial direction R, and thus, the surface area of the sidewalls 199 defining the blind warming openings 190 is increased (e.g., as opposed to the sidewalls of blind warming openings oriented parallel with the radial direction R), which ultimately increases the rate of heat transfer or thermal response of the forward end 112 .
- the plurality of blind warming openings 190 are oriented at an angle ⁇ with respect to the radial direction R. More particularly, for this embodiment, the outer surface openings 196 are angled at a negative forty-five degree ( ⁇ 45°) angle with respect to the radial direction R and the inner surface openings 198 are angled at a forty-five degree (45°) angle with respect to the radial direction R. Thus, the outer surface openings 196 are angled opposite the inner surface openings 198 with respect to the radial direction R.
- the axial spacing of the blind warming openings 190 can be positioned so as to provide efficient warming of the forward end 112 along its radial thickness and axial length without compromising the structural integrity of the forward end 112 .
- angling the blind warming openings 190 increases the surface area of the sidewalls 199 defining the openings, thereby increasing thermal response of the forward end 112 .
- the blind warming openings 190 can be oriented in axial rows as shown in FIG. 4 along the circumferential direction C along the forward end 112 and between and/or about the mounting openings 174 .
- the outer surface openings 196 are angled at a negative forty-five degree ( ⁇ 45°) angle with respect to the radial direction R and the inner surface openings 198 are likewise angled at a negative forty-five degree ( ⁇ 45°) angle with respect to the radial direction R.
- the outer surface openings 196 are not angled opposite the inner surface openings 198 with respect to the radial direction R.
- this may increase the rate of heat transfer of the forward end 112 .
- the warming openings 190 are angled at a sixty degree (60°) angle (positive or negative) with respect to the radial direction R, thereby further increasing the surface area of the sidewalls 199 defining the blind warming openings 190 .
- the blind warming openings 190 are angled at a seventy-five degree (75°) (positive or negative) angle with respect to the radial direction R.
- the blind warming openings 190 are oriented at an angle that is greater than forty-five degrees (45°) (positive or negative) with respect to the radial direction R.
- the blind warming openings 190 are oriented approximately along the radial direction R. By orienting the blind warming openings 190 along approximately along the radial direction R, the openings may be easier to machine into the forward end 112 of the outer liner 108 and may be less impactful to the structural integrity of the forward end 112 .
- the temperature of the generated combustion gases increases.
- the high temperature combustion gases scrub along the outer and inner liners, which rapidly heats the liners.
- the forward ends of the liners do not heat up as quickly.
- the forward ends of the liners thermally lag the other portions of the liners. Consequently, the liners may experience bending stress and strain due to the thermal lag of the forward ends.
- the blind warmings openings allow for a warming airflow to flow therein to warm the forward ends of the inner and outer liners. In this way, the stress and strain on the liners can be reduced, which as noted above, may improve the durability of the liners.
- a warming airflow WA may flow into the blind warming openings 190 as follows. As shown, during operation of turbofan engine 10 ( FIG. 1 ), the warming airflow WA flows into the slot 122 defined by the outer dome 118 . More particularly, the warming air WA flows between the yoke 160 and the outer surface 178 of the outer liner 108 . In this example, the warming airflow WA is compressor discharge air (P3 air). As the warming airflow WA travels forward along the axial direction A deeper into the slot 122 , some of the warming airflow WA flows into the blind warming openings 190 , and more particularly the outer surface openings 196 , as shown. The warming airflow WA flows into the outer surface openings 196 and convection heat transfer occurs.
- P3 air compressor discharge air
- the warming airflow WA flows into the outer surface openings 196 and exchanges heat with the sidewalls 199 of the openings, thus effectively warming the forward end 112 of the outer liner 108 . Further, some of the warming airflow WA that enters slot 122 flows past the outer surface openings 196 and flows radially inward through the axial gap G A ( FIG. 3 ) between the forward surface 121 of the outer dome 118 and the radial interface surface 188 of the forward end 112 . Thereafter, the warming airflow WA flows aft along the axial direction A between the inner surface 182 of the forward end 112 and the base plate 158 of the outer dome 118 . Some of the warming airflow WA flows into the inner surface openings 198 and convection heat transfer occurs.
- the warming airflow WA flows into the inner surface openings 198 and exchanges heat with the sidewalls 199 of the openings, thus effectively warming the forward end 112 of the outer liner 108 .
- the warming airflow WA flows aft along the axial direction A through the radial gap G R ( FIG. 3 ) and into the combustion chamber 114 .
- the forward end 112 of the outer liner 108 is actively warmed. Accordingly, during transient bursts of the engine, the transient response rate of the forward end 112 is improved thereby reducing the bending stress and strain on the liner.
- the forward end 106 of the inner liner 102 may be formed in the same or substantially the same manner as the forward end 112 of the outer liner 108 as described above. Further, it will be appreciated that the forward end 106 of the inner liner 102 may be attached to the inner dome section 116 in the same or substantially the same manner that the forward end 112 of the outer liner 108 is attached to the outer dome section 118 .
- FIG. 5 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118 . Further, FIG. 5 depicts the forward end 112 of outer liner 108 defining blind warming opening 190 , which in the depicted embodiment of FIG. 5 is a blind warming slit 200 . Like the blind warming openings depicted in the embodiment of FIG. 4 , the warming slit 200 defined by the forward end 112 increases the thermal response of the forward end 112 of the outer liner 108 . Although only one (1) warming slit 200 is shown in FIG. 5 , in some embodiments the forward end 112 may define a plurality of warming slits 200 .
- the forward end 112 of the outer liner 108 may define a plurality of warming slits 200 spaced apart from one another along the circumferential direction C.
- the forward end 112 may define more than one warming slit 200 along the radial direction R.
- the forward end 106 of the inner liner 102 can likewise define blind warming slits 200 .
- the diameter of the blind warming slits 200 may range between or about between 0.020 and 0.10 inches (0.508-2.54 mm).
- the diameter of the mounting openings 174 may range between or about between 0.400 and 0.800 inches (10.16-20.32 mm). As such, in such embodiments, the mounting openings 174 have larger diameters than the blind warming slits 200 .
- the warming slit 200 extends between an opening end 202 and a terminus end 204 along the axial direction A.
- the radial interface surface 188 defines the opening end 202 of the warming slit 200 .
- the warming slit 200 extends along the axial direction A aft of the mounting opening 174 . In this way, during operation of turbofan engine 10 ( FIG. 1 ), the forward end 112 is warmed by warming airflow WA from the radial interface surface 188 to a point along the axial length of the forward end 112 that is aft of the mounting assembly 144 that couples the forward end 112 of the outer liner 108 with the outer dome 118 .
- the warming slit 200 is defined by the forward end 112 along the axial direction A from the radial interface surface 188 to a position that is aligned with or aft of an edge 143 along the axial direction A.
- the edge 143 is a radial outermost point of a trailing edge 145 of heat shield 142 .
- the heat shield 142 generally facilitates the flow of combustion gases 66 downstream of the forward end 112 of the outer liner 108 ( FIG. 2 )
- the forward end 112 is susceptible to thermal lag from its radial interface surface 188 to a point along the axial direction A that is protected by the heat shield 142 .
- the warming slit 200 along this length effectively warms the forward end 112 during transient operation of the turbofan engine 10 ( FIG. 1 ), and thus, bending stress and strain on the liner may be reduced.
- the warming slit 200 is defined by the forward end 112 along the axial direction A from the radial interface surface 188 to a position that is aligned with or aft of an edge 147 along the axial direction A.
- the edge 147 is a radial outermost point of a trailing edge 149 of outer dome 118 .
- the outer dome 118 generally shields the forward end 112 of the outer liner 108 from combustion gases 66
- the forward end 112 is susceptible to thermal lag from its radial interface surface 188 to a point along the axial direction A that is shielded by the outer dome 118 .
- the warming slit 200 along this length effectively warms the forward end 112 during transient operation of the turbofan engine 10 ( FIG. 1 ), and thus, bending stress and strain on the liner may be reduced.
- warming airflow WA may flow into the blind warming openings 190 , which in this embodiment is warming slit 200 , as follows. As shown, during operation of turbofan engine 10 ( FIG. 1 ), the warming airflow WA flows into the slot 122 defined by the outer dome 118 . More particularly, the warming air WA flows between the yoke 160 and the outer surface 178 of the outer liner 108 . In this example, the warming airflow WA is compressor discharge air (P3 air). After the warming airflow WA flows into the slot 122 , the warming airflow WA flows around the most forward portion of the forward end 112 and flows into axial gap G A .
- P3 air compressor discharge air
- the warming airflow WA flows through the axial gap G A , some of the warming airflow WA flows into the opening end 202 of the warming slit 200 .
- the warming airflow WA then travels forward along the axial direction A toward the terminus end 204 of the warming slit 200 .
- convection heat transfer occurs. That is, the warming airflow WA flowing through warming slit 200 exchanges heat with the walls 206 defining the warming slit 200 , thus effectively warming the forward end 112 of the outer liner 108 .
- the warming airflow WA flows aft along the axial direction A between the inner surface 182 of the outer liner 108 and the base plate 158 and through the radial gap G R and into the combustion chamber 114 .
- the forward end 112 of the outer liner 108 is actively warmed. Accordingly, during transient bursts of the engine, the transient response rate of the forward end 112 may be improved, which may reduce the bending stress and strain on the liner.
- forward end 112 of outer liner 108 defines outer surface openings 196 ( FIG. 4 ), inner surface openings 198 ( FIG. 4 ), and warming slits 200 ( FIG. 5 ).
- the forward end 106 of the inner liner 102 can likewise include such configurations.
- FIG. 6 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118 .
- FIG. 6 further depicts the outer dome 118 defining a plurality of impingement openings 210 .
- the yolk 160 of the outer dome 118 defines the impingement openings 210 .
- the yolk 160 extends between an outer surface 161 and an inner surface 163 along the radial direction R, and for this embodiment, the impingement openings 210 extend from the outer surface 161 to the inner surface 163 of the yolk 160 .
- the impingement openings 210 are through holes that extend through the yolk 160 .
- the impingement openings 210 each extend between an outer end 212 and an inner end 214 .
- the outer ends 212 of the impingement openings 210 are defined by the outer surface 161 of the yolk 160 and the inner ends 214 of the impingement openings 210 are defined by the inner surface 163 of the yolk 160 .
- the outer end 212 of each impingement opening 210 has an outer diameter D O and the inner end 214 of each impingement opening 210 has an inner diameter D I .
- the outer diameters D O of the impingement openings 210 are greater than the inner diameters D I .
- the impingement openings 210 taper as they extend radially inward from the outer surface 161 to the inner surface 163 of the yolk 160 .
- warming airflow WA e.g., compressor discharge air
- the warming airflow WA flows into the impingement openings 210 .
- the warming airflow WA enters the impingement openings 210 through their respective outer ends 212 .
- the pressure of the warming airflow increases due to the converging geometry of the impingement openings 210 .
- the flow rate of the warming airflow WA rapidly increases thereby creating a high velocity or jet stream of warming airflow WA.
- the inner diameters D I of the impingement openings 210 are greater than the outer diameters D O .
- the impingement openings 210 taper as they extend radially outward from the outer surface 161 to the inner surface 163 of the yolk 160 .
- the warming airflow WA flowing radially outward of the yolk 160 can more easily flow into the impingement openings 210 , among other benefits.
- the inner diameters D I of the impingement openings 210 are the same as the outer diameters D O . That is, the impingement openings 210 do not taper as they extend along the radial direction R.
- the impingement openings 210 are angled with respect to the radial direction R.
- the inner ends 214 of the impingement openings 210 may be positioned aft of their respective outer ends 212 . In this way, the warming airflow WA may flow into the impingement openings 210 more easily.
- FIG. 7 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118 .
- FIG. 7 further depicts the outer dome 118 defining impingement openings 210 .
- FIG. 7 also depicts an impingement jacket 220 attached to the outer dome 118 .
- the impingement jacket 220 extends between a forward end 222 and an aft end (not shown) generally along the axial direction A.
- the impingement jacket 220 spans the axial length of the outer liner 108 (i.e., from the forward end 112 to the aft end 110 of the outer liner 108 ; see FIG. 2 ).
- the forward end 222 of the impingement jacket 220 is attached to the yolk 160 of the outer dome 118 .
- the impingement jacket 220 may be attached to the yolk 160 in any suitable manner.
- the impingement jacket 220 may extend annularly about the circumferential direction C or may extend along the circumferential direction C in segments.
- the impingement jacket 220 defines a plurality of impingement openings 224 .
- the impingement openings 224 may be configured in any suitable manner as noted above with respect to the impingement openings 210 defined by the yolk 160 of the outer dome 118 .
- the diameter of the blind warming openings 190 may range between or about between 0.020 and 0.080 inches (0.508-2.032 mm). Accordingly, during operation of turbofan engine 10 ( FIG. 1 ), warming airflow WA may flow through the impingement openings 224 defined by the impingement jacket 220 and impinge on the outer surface 178 of the forward end 112 of the outer liner 108 .
- FIG. 8 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118 .
- FIG. 8 depicts metal grommet 172 of mounting assembly 144 positioned within mounting opening 174 and defining a plurality of warming passages 230 .
- the warming passages 230 may be drilled into or otherwise machined into the grommet 172 .
- the warming passages 230 defined by grommet 172 each extend between an outer end 232 and an inner end 234 .
- the warming passages 230 defined by grommet 172 extend along the radial direction R.
- the outer ends 232 of the warming passages 230 are defined by the outer collar 176 of the grommet 172 , while the inner ends 234 of the warming passages 230 are defined by the inner collar 180 of the grommet 172 . Further, as the warming passages 230 extend through the grommet 172 , the warming passages 230 also extend through the body 184 of the grommet 172 . Notably, as shown in FIG. 8 , the warming passages 230 extend through the body 184 of the grommet 172 such that the warming passages 230 extend along a radial face 113 of the forward end 112 of the outer liner 108 . In this way, during operation of the turbofan engine 10 ( FIG. 1 ), warming airflow WA may flow through the warming passages 230 and warm the radial face 113 of the forward 112 , thereby actively warming the forward end 112 .
- FIG. 9 provides a cross sectional view of the grommet 172 of FIG. 8 taken along line 9 - 9 of FIG. 8 .
- the bushing 164 is generally cylindrical in shape and is positioned around the shank 168 of the pin 162 .
- the bushing 164 and pin 162 are positioned within and extend through a grommet opening 173 defined by grommet 172 .
- the grommet 172 is positioned in the mounting opening 174 defined by the forward end 112 of the outer liner 108 .
- the warming passages 230 are circumferentially spaced about the perimeter 185 of the body 184 of the grommet 172 .
- Each warming passage 230 extends through the perimeter 185 of the body 184 as a cutout and has a semi-circle shaped radial cross section and is, as noted above, positioned or defined by grommet 172 such that the warming passages 230 are defined in part by the grommet 172 and defined in part by the forward end 112 of the outer liner 108 .
- This may allow for warming airflow WA to pass therethrough and actively warm the radial face 113 and the forward end 112 more generally.
- the transient thermal response rate of the forward end 112 can be increased, and accordingly, the bending stress and strain on the outer liner 108 may be reduced.
- the grommet 172 may be sized such that a gap is defined between the grommet 172 and the radial face 113 of the forward end 112 . In this way, during operation of the turbofan engine 10 (FIG. 1 ), warming airflow WA may flow through the gap and warm the radial face 113 , thereby actively warming the forward end 112 of the outer liner 108 .
- FIG. 10 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118 . Further, FIG. 10 depicts the forward end 112 having enhanced surfaces.
- the outer surface 178 of the outer liner 108 has a heat transfer coefficient (HTC) enhanced surface 240 and the inner surface 182 of the outer liner 108 likewise has an HTC enhanced surface 240 .
- HTC heat transfer coefficient
- the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may be a rough bond coating.
- the coating may a suitable ceramic thermal and environmental barrier coating (TEBC) for CMC components.
- TEBC ceramic thermal and environmental barrier coating
- the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may have an undulating surface as shown in FIG. 10 .
- the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may be a grit blasted coating.
- the outer and/or inner surfaces 178 , 182 of the forward end 112 may be blasted with a coarse grit.
- the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may have a plurality of bumps having a height of about ten (10) millimeters.
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Abstract
Description
- The present subject matter relates generally to gas turbine engines, and more particularly to combustor assemblies for gas turbine engines.
- 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 the ability of 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, inner and outer liners of gas turbine engines are more commonly being formed of CMC materials.
- During normal operation, it is common for gas turbine engines to be required to rapidly increase thrust. During such rapid power increases or transient state conditions, combustion gases are generated within a combustion chamber defined by inner and outer liners. As the combustion gases flow downstream through the combustion chamber, the combustion gases scrub along the liners, causing the liners to rapidly heat up. However, the forward ends of the liners, or the portions of the liners that attach with dome sections, typically do not heat up as quickly as the rest of their respective liners. The thermal lag at the forward ends of the liners may cause undesirable bending stress and strain on the liners. As gas turbine engines may undergo many rapid power increases over many engine cycles, such repeated stress and strain on the liners can negatively impact their durability.
- Accordingly, a combustor assembly of a gas turbine engine that includes features that reduce the stress and strain on combustion liners of the combustor assembly during rapid power increases of the gas turbine engine would be useful.
- 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 defining an axial direction, a radial direction, and a circumferential direction is provided. The combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end of the liner defines a plurality of mounting openings spaced along the circumferential direction. The forward end of the liner also defines one or more blind warming openings.
- In another exemplary aspect of the present disclosure, a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction is provided. The combustor assembly includes a dome comprising a yolk and a base plate spaced outward from the yolk along the radial direction, the yolk and the base plate defining a slot. Further, the combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot. The yolk defines a plurality of impingement openings extending therethrough such that a warming airflow impinges on the forward end of the liner during operation of the gas turbine engine.
- In yet another exemplary aspect of the present disclosure, a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction is provided. The combustor assembly includes a dome defining a slot and a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot of the dome, the forward end of the liner defining a plurality of mounting openings spaced along the circumferential direction. Further, the combustor assembly includes a plurality of mounting assemblies for coupling the forward end of the liner with the dome, each of the plurality of mounting assemblies including: a pin extending through one of the plurality of mounting openings and a grommet positioned within one of the plurality of mounting openings and defining a grommet opening, the pin received within the grommet opening, and wherein the grommet defines one or more warming passages.
- 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 provides a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter; -
FIG. 2 provides a schematic, cross-sectional view of a combustor assembly in accordance with an exemplary embodiment of the present disclosure; -
FIG. 3 provides 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 dome section; -
FIG. 4 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and a plurality of blind warming openings defined by the forward end in accordance with an exemplary embodiment of the present disclosure; -
FIG. 5 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and a warming slit defined by the forward end in accordance with an exemplary embodiment of the present disclosure; -
FIG. 6 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and impingement openings defined by the outer dome section in accordance with an exemplary embodiment of the present disclosure; -
FIG. 7 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and depicts an impingement jacket attached to the outer dome section in accordance with an exemplary embodiment of the present disclosure; -
FIG. 8 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and warming passages defined by a grommet in accordance with an exemplary embodiment of the present disclosure; -
FIG. 9 provides a cross sectional view of the grommet ofFIG. 8 taken on line 9-9 ofFIG. 8 ; and -
FIG. 10 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and further depicting the forward end of the outer liner having enhanced surfaces in accordance with an exemplary embodiment of the present disclosure. - 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 “forward” and “aft” refer to relative positions within a gas turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust. 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. It should be appreciated, that as used herein, terms of approximation, such as “about” and “approximately,” refer to being within a ten percent (10%) margin of error.
- Exemplary aspects of the present disclosure are directed to a combustor assembly for gas turbine engines. In one exemplary aspect, the combustor assembly includes a dome defining a slot and a liner that at least partially defines a combustion chamber. The liner extends between an aft end and a forward end. At least a portion of the forward end is received within the slot of the dome. In one aspect, the forward end of the liner defines a plurality of blind warming openings that allow a warming airflow to actively warm the forward end during transient operation of the engine. In another aspect, the dome defines a plurality of impingement openings that allow a warming airflow to impinge on the forward end of the liner to actively warm the forward end during transient operation of the engine. In another aspect, a grommet coupling a pin with the liner defines a plurality of warming passages that allow a warming airflow to flow therethrough to warm the forward end of the liner during transient engine operation. In a further aspect, one or more surfaces of the forward end of the liner have enhanced surfaces that increase the thermal response of the forward end. By warming the forward end of the liner, the stress and strain on the liner during transient operating conditions may be reduced, particularly during transient engine power increases or bursts. Moreover, by reducing the stress and strain on the liner, the durability of the liner or liners may be improved.
-
FIG. 1 provides 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-bypass turbofan jet engine 10, referred to herein as “turbofan engine 10.” As shown inFIG. 1 , the turbofan engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). In general, the turbofan 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 theHP turbine 28 to theHP compressor 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 by rotatable spinner orfront 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 of
air 58 enters the turbofan 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 the turbofan 10, also providing propulsive thrust. TheHP turbine 28, theLP turbine 30, and the jetexhaust nozzle section 32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through thecore turbine engine 16. - It should be appreciated that the exemplary turbofan engine 10 depicted in
FIG. 1 is by way of example only, and that in other exemplary embodiments, the turbofan engine 10 may have any suitable configuration. For example, the present disclosure matter may be suitable for use with or in turboprops, turboshafts, turbojets, reverse-flow engines, industrial and marine gas turbine engines, and/or auxiliary power units. -
FIG. 2 provides a close-up cross-sectional view of acombustor assembly 100 in accordance with an exemplary embodiment of the present disclosure. For example, thecombustor assembly 100 ofFIG. 2 may be positioned in thecombustion section 26 of the exemplary turbofan engine 10 ofFIG. 1 . More particularly,FIG. 2 provides a side, cross-sectional view of theexemplary combustor assembly 100 ofFIG. 2 . - As shown, the
combustor assembly 100 includes aninner liner 102 extending between anaft end 104 and aforward end 106 generally along the axial direction A, as well as anouter liner 108 also extending between anaft end 110 and aforward end 112 generally along the axial direction A. The inner andouter liners combustion chamber 114 therebetween. The inner andouter liners inner dome section 116 attached to theforward end 106 of theinner liner 102 and anouter dome section 118 attached to theforward end 112 of theouter liner 108. The inner andouter dome sections FIG. 3 , the inner andouter dome sections forward surface 121 at least partially defining aslot 122 for receipt of theforward end 106 of theinner liner 102, and theforward end 112 of theouter liner 108, respectively. - The
combustor assembly 100 further includes a plurality offuel air mixers 124 spaced along a circumferential direction C and positioned at least partially within the annular dome. More particularly, the plurality offuel air mixers 124 are disposed at least partially between theouter dome section 118 and theinner dome section 116 along the radial direction R. Compressed air from the compressor section of the turbofan engine 10 flows into or through thefuel air mixers 124, where the compressed air is mixed with fuel and ignited to create the combustion gases 66 (FIG. 1 ) within thecombustion chamber 114. The inner andouter dome sections fuel air mixers 124. For example, theouter dome section 118 includes anouter cowl 126 at aforward end 128 and theinner dome section 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 the compressor section into or through one or more of thefuel air mixers 124. - Moreover, the inner and
outer dome sections combustor assembly 100 within the turbofan engine 10 (FIG. 1 ). For example, theouter dome section 118 includes anattachment extension 134 configured to be mounted to anouter combustor casing 136 and theinner dome section 116 includes asimilar attachment extension 138 configured to attach to anannular support member 140 within the turbofan engine 10. In certain exemplary embodiments, theinner dome section 116 may be formed integrally as a single annular component, and similarly, theouter dome section 118 may also be formed integrally as a single annular component. It should be appreciated, however, that in other exemplary embodiments, theinner dome section 116 and/or theouter dome section 118 may alternatively be formed by one or more components being joined in any suitable manner. For example, with reference to theouter dome section 118, in certain exemplary embodiments, theouter cowl 126 may be formed separately from theouter dome section 118 and attached to theforward end 128 of theouter dome section 118 using, e.g., a welding process. Similarly, theattachment extension 134 may also be formed separately from theouter dome section 118 and attached to theforward end 128 of theouter dome section 118 using, e.g., a welding process. Additionally or alternatively, theinner dome section 116 may have a similar configuration. - With reference still to
FIG. 2 , theexemplary combustor assembly 100 further includes aheat shield 142 positioned around thefuel air mixer 124 depicted. For this embodiment, theexemplary heat shield 142 is attached to and extends between theouter dome section 118 and theinner dome section 116. Theheat shield 142 is configured to protect certain components of the turbofan engine 10 (FIG. 1 ) from the relatively extreme temperatures of thecombustion chamber 114. - For the embodiment depicted, the
inner liner 102 and theouter liner 108 are each formed 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 annular dome, including the
inner dome section 116 andouter dome section 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.). - Referring still to
FIG. 2 , as noted above, the combustion gases 66 (FIG. 1 ) flow from thecombustion chamber 114 into and through the turbine section of the turbofan engine 10 (FIG. 1 ) 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 one (1)turbine nozzle 156 is depicted schematically inFIG. 2 positioned aft of thecombustor assembly 100. -
FIG. 3 provides a close up, schematic, cross-sectional view of an attachment point where theforward end 112 of theouter liner 108 is mounted to theouter dome section 118 within theslot 122 of theouter dome section 118. To allow for a relative thermal expansion between theouter liner 108 and theouter dome section 118, as well as between theinner liner 102 and the inner dome section 116 (FIG. 2 ), a plurality of mountingassemblies 144 are used to attach theouter liner 108 to theouter dome section 118 and theinner liner 102 to theinner dome section 116. More particularly, the mountingassemblies 144 attach theforward end 112 of theouter liner 108 to theouter dome section 118 within theslot 122 of theouter dome section 118 as shown inFIGS. 2 and 3 and theforward end 106 of theinner liner 102 to theinner dome section 116 within theslot 122 of the inner dome section 116 (FIG. 2 ). Theslots 122 are defined by their respective domes. Moreover, theslots 122 receive the forward ends 106, 112 of the inner andouter liners - Referring particularly to the
forward end 112 of theouter liner 108 and theouter dome section 118 depicted inFIG. 3 , theouter dome section 118 includes abase plate 158 and ayolk 160. Thebase plate 158 and theyolk 160 are spaced along the radial direction R. Thebase plate 158 and theyolk 160 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 (see alsoFIG. 2 ). Notably, theslot 122 is defined between thebase plate 158 and theyolk 160. Theslot 122 is further defined by theforward surface 121. Further, in certain exemplary embodiments, theyolk 160 may extend circumferentially with theouter dome section 118, tracking thebase plate 158. With such a configuration, theslot 122 may be considered an annular slot. However, in other embodiments, theyolk 160 may include a plurality of circumferentially spaced tabs, each of the individual tabs of theyolk 160 defining individual segmented portions of theslot 122 with thebase plate 158. - The exemplary mounting
assembly 144 depicted includes theyolk 160 of theouter dome section 118 and thebase plate 158 of theouter dome section 118. Moreover, the mountingassembly 144 includes apin 162 and abushing 164. Thepin 162 includes ahead 166 and ashank 168. Theshank 168 extends through theyolk 160, theforward end 112 of the outer liner 108 (positioned in slot 122), and thebase plate 158. Anut 170 is attached to a distal end of theshank 168 of thepin 162. In certain exemplary embodiments, thepin 162 may be configured as a bolt and thenut 170 may be rotatably engaged with a threaded portion of the pin 162 (at, e.g., the distal end of the shank 168) for tightening the mountingassembly 144. Alternatively, however, in other exemplary embodiments thepin 162 andnut 170 may have any other suitable configurations. In other exemplary embodiments, for instance, thepin 162 may include ashank 168 defining a substantially smooth cylindrical shape and thenut 170 may be configured as a clip. - Additionally, the
bushing 164 is generally cylindrical in shape and is positioned around theshank 168 of thepin 162 within theslot 122. For the embodiment depicted, thebushing 164 is pressed between theyolk 160 and thebase plate 158 by tightening thenut 170 on thepin 162. Moreover, for the embodiment depicted, the mountingassembly 144 includes ametal grommet 172 positioned around thebushing 164 andpin 162. Thegrommet 172 is positioned in a mountingopening 174 defined by theforward end 112 of theouter liner 108. Thegrommet 172 includes anouter collar 176 positioned adjacent to anouter surface 178 of theouter liner 108 and aninner collar 180 positioned adjacent to aninner surface 182 of theouter liner 108. Thegrommet 172 additionally includes abody 184. Themetal grommet 172 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 section 118. - It should be appreciated, however, that although the
forward end 112 of theouter liner 108 is attached to theouter dome section 118 using the exemplary mountingassembly 144 depicted and described herein, in other embodiments of the present disclosure, the mountingassembly 144 may have other suitable configurations, and further still in other embodiments, any other suitable attachment assembly may be used. - Referring still to
FIG. 3 , theforward end 112 of theouter liner 108 depicted further includes anaxial interface surface 186 and aradial interface surface 188. Theaxial interface surface 186 is configured as a portion of theforward end 112 of theouter liner 108 facing thebase plate 158 of theouter dome section 118, or more particularly, facing theinner surface 120 of theouter dome section 118. Theradial interface surface 188 is configured as a portion of theforward end 112 of theouter liner 108 facing theforward surface 121 of theouter dome section 118. For the embodiment depicted, theaxial interface surface 186 andinner surface 120 each extend in a direction parallel to the axial direction A, and theradial interface surface 188 andforward surface 121 each extend in a direction parallel to the radial direction R. - Moreover, as further shown in
FIG. 3 , theaxial interface surface 186 defines a radial gap GR with theinner surface 120 of theouter dome section 118 and theradial interface surface 188 defines an axial gap GA with theforward surface 121 of theouter dome section 118. Thecombustor assembly 100 may be designed such that the radial and axial gaps GR, GA allow for only a predetermined amount of airflow therethrough into thecombustion chamber 114. Notably, allowing such a flow of air during operating conditions of thecombustor assembly 100 may ensure relatively hot combustion gases within thecombustion chamber 114 do not flow into and/or through theslot 122 of theouter dome section 118, potentially damaging certain components of thecombustor assembly 100. - In addition to the airflow through the radial and axial gaps GR, GA, in some exemplary embodiments as will be explained more fully below, additional airflow can be provided to warm the
forward end 112 of the outer liner 108 (as well as theforward end 106 of theinner liner 102 depicted inFIG. 2 ) to improve the thermal response (e.g., reduce the thermal lag) of the forward ends 112, 106 of the outer andinner liners FIG. 1 ), and particularly during rapid power increases or bursts of the turbofan engine 10. In this way, the stress and strain on the outer andinner liners -
FIG. 4 provides a close up, cross-sectional view of one exemplary embodiment ofcombustor assembly 100 depicting forward end 112 ofouter liner 108 attached toouter dome 118. Further,FIG. 4 depicts theforward end 112 of theouter liner 108 defining a plurality ofblind warming openings 190. The diameter of theblind warming openings 190 may range between or about between 0.020 and 0.080 inches (0.508-2.032 mm). The diameter of the mountingopenings 174 may range between or about between 0.400 and 0.800 inches (10.16-20.32 mm). Thus, the mountingopenings 174 have larger diameters than theblind warming openings 190. Theblind warming openings 190 defined by theforward end 112 are operatively configured to increase the thermal response of theforward end 112 of theouter liner 108. As such, as will be explained more fully below, the stress and strain on theouter liner 108 can be reduced, thereby improving the durability of theouter liner 108. Theforward end 106 of theinner liner 102 can likewise defineblind warming openings 190, and thus, the durability of theinner liner 102 can likewise be improved. - As shown in
FIG. 4 , each of theblind warming openings 190 extend between anopening end 192 and aterminus end 194. The opening ends 192 of theblind warming openings 190 allow fluid (e.g., a warming airflow WA) to flow into and out of their respective openings and the terminus ends 194 are the dead ends or terminating points of theblind warming openings 190. For the depicted embodiment ofFIG. 4 , theblind warming openings 190 extend along the radial direction R from one of theouter surface 178 and theinner surface 182 and terminate prior to the opposing surface. In this way, the warming openings are blind openings (i.e., they do not extend through theforward end 112 of the outer liner 108). More particularly, as shown for the depicted embodiment ofFIG. 4 , one or more of theblind warming openings 190 extend along the radial direction R from theouter surface 178 and terminate prior to the opposinginner surface 182 and one or more of theblind warming openings 190 extend along the radial direction R from theinner surface 182 and terminate prior to the opposingouter surface 178. Thus, theblind warming openings 190 includeouter surface openings 196 that extend along the radial direction R from theouter surface 178 and terminate prior to theinner surface 182 and theblind warming openings 190 includeinner surface openings 198 that extend along the radial direction R from theinner surface 182 and terminate prior to theouter surface 178. The opening ends 192 of theouter surface openings 196 are defined by theouter surface 178 of theforward end 112 and the terminus ends 194 of theouter surface openings 196 are defined by theforward end 112 at some point along the radial thickness of theforward end 112. In contrast, the opening ends 192 of theinner surface openings 198 are defined by theinner surface 182 of theforward end 112 and the terminus ends 194 of theinner surface openings 198 are defined by theforward end 112 at some point along the radial thickness of theforward end 112. - As further shown in the circumferential cross section of
FIG. 4 , theforward end 112 defines a row ofblind warming openings 190. More particularly, the row ofblind warming openings 190 includesouter surface openings 196 that alternate with theinner surface openings 198 along the axial direction A. By alternating theouter surface openings 196 with theinner surface openings 198 along the axial direction A, theouter surface openings 196 and theinner surface openings 198 can be axially positioned near one another without compromising the structural integrity of theforward end 112 of theouter liner 108. In this way, efficient warming of theforward end 112 along its radial thickness and axial length can be achieved without, as noted previously, compromising the structural integrity of theforward end 112. - As further depicted in
FIG. 4 , a midline is defined midway between the outer andinner surfaces forward end 112 of theouter liner 108 along the radial direction R. As shown, for this embodiment, theblind warming openings 190 extend along the radial direction R from one of the outer andinner surfaces forward end 112 past the midline M. More particularly, as shown, theouter surface openings 196 extend from theouter surface 178 past the midline M and theinner surface openings 198 extend from theinner surface 182 past the midline M. More particularly still, the opening ends 192 of theouter surface openings 196 extend from theouter surface 178 and their terminus ends 194 extend past the midline M. Likewise, the opening ends 192 of theinner surface openings 198 extend from theinner surface 182 and their terminus ends 194 extend past the midline M. As theblind warming openings 190 extend along the radial direction R past the midline M in the depicted embodiment ofFIG. 4 , it is ensured that theforward end 112 is warmed along its radial thickness. - As further depicted in the exemplary embodiment of
FIG. 4 , theblind warming openings 190 are angled with respect to the radial direction R. For this embodiment, for instance, the opening ends 192 of theouter surface openings 196 and theinner surface openings 198 are positioned aft of their respective terminus ends 194 along the axial direction A. Stated differently, the opening ends 192 of theblind warming openings 190 are not aligned with their respective terminus ends 194 along the radial direction R. In this manner, theblind warming openings 190 are angled with respect to the radial direction R, and thus, the surface area of thesidewalls 199 defining theblind warming openings 190 is increased (e.g., as opposed to the sidewalls of blind warming openings oriented parallel with the radial direction R), which ultimately increases the rate of heat transfer or thermal response of theforward end 112. - Additionally, for the depicted embodiment of
FIG. 4 , the plurality ofblind warming openings 190 are oriented at an angle θ with respect to the radial direction R. More particularly, for this embodiment, theouter surface openings 196 are angled at a negative forty-five degree (−45°) angle with respect to the radial direction R and theinner surface openings 198 are angled at a forty-five degree (45°) angle with respect to the radial direction R. Thus, theouter surface openings 196 are angled opposite theinner surface openings 198 with respect to the radial direction R. By angling the outer andinner surface openings blind warming openings 190 can be positioned so as to provide efficient warming of theforward end 112 along its radial thickness and axial length without compromising the structural integrity of theforward end 112. Moreover, as noted above, angling theblind warming openings 190 increases the surface area of thesidewalls 199 defining the openings, thereby increasing thermal response of theforward end 112. Theblind warming openings 190 can be oriented in axial rows as shown inFIG. 4 along the circumferential direction C along theforward end 112 and between and/or about the mountingopenings 174. - In other exemplary embodiments, the
outer surface openings 196 are angled at a negative forty-five degree (−45°) angle with respect to the radial direction R and theinner surface openings 198 are likewise angled at a negative forty-five degree (−45°) angle with respect to the radial direction R. Thus, theouter surface openings 196 are not angled opposite theinner surface openings 198 with respect to the radial direction R. By angling theouter surface openings 196 at a negative angle with respect to the radial direction R, as warming airflow WA flows into theslot 122 between theyolk 160 and theforward end 112 of theouter liner 108 the warming airflow WA flows more directly into theouter surface openings 196. Thus, this may increase the rate of heat transfer of theforward end 112. Similarly, by angling theinner surface openings 198 at a negative angle with respect to the radial direction R, the warming airflow WA flowing into theslot 122 between theyolk 160 and theforward end 112, through the axial gap GA (FIG. 3 ), and then aft between theinner surface 182 of theliner 108 and thebase plate 158 flows more directly into theinner surface openings 198. This may increase the rate of heat transfer of theforward end 112. - In yet other exemplary embodiments, the warming
openings 190 are angled at a sixty degree (60°) angle (positive or negative) with respect to the radial direction R, thereby further increasing the surface area of thesidewalls 199 defining theblind warming openings 190. In further embodiments, theblind warming openings 190 are angled at a seventy-five degree (75°) (positive or negative) angle with respect to the radial direction R. In some embodiments, theblind warming openings 190 are oriented at an angle that is greater than forty-five degrees (45°) (positive or negative) with respect to the radial direction R. It yet further embodiments, theblind warming openings 190 are oriented approximately along the radial direction R. By orienting theblind warming openings 190 along approximately along the radial direction R, the openings may be easier to machine into theforward end 112 of theouter liner 108 and may be less impactful to the structural integrity of theforward end 112. - During transient operation of the turbofan engine, and particularly during a rapid power increase or burst, the temperature of the generated combustion gases increases. The high temperature combustion gases scrub along the outer and inner liners, which rapidly heats the liners. However, as noted previously, the forward ends of the liners do not heat up as quickly. Stated differently, the forward ends of the liners thermally lag the other portions of the liners. Consequently, the liners may experience bending stress and strain due to the thermal lag of the forward ends. To reduce the thermal lag and thus the stress and strain on the liners, the blind warmings openings allow for a warming airflow to flow therein to warm the forward ends of the inner and outer liners. In this way, the stress and strain on the liners can be reduced, which as noted above, may improve the durability of the liners.
- With reference to
FIG. 4 , a warming airflow WA may flow into theblind warming openings 190 as follows. As shown, during operation of turbofan engine 10 (FIG. 1 ), the warming airflow WA flows into theslot 122 defined by theouter dome 118. More particularly, the warming air WA flows between theyoke 160 and theouter surface 178 of theouter liner 108. In this example, the warming airflow WA is compressor discharge air (P3 air). As the warming airflow WA travels forward along the axial direction A deeper into theslot 122, some of the warming airflow WA flows into theblind warming openings 190, and more particularly theouter surface openings 196, as shown. The warming airflow WA flows into theouter surface openings 196 and convection heat transfer occurs. More specifically, the warming airflow WA flows into theouter surface openings 196 and exchanges heat with thesidewalls 199 of the openings, thus effectively warming theforward end 112 of theouter liner 108. Further, some of the warming airflow WA that entersslot 122 flows past theouter surface openings 196 and flows radially inward through the axial gap GA (FIG. 3 ) between theforward surface 121 of theouter dome 118 and theradial interface surface 188 of theforward end 112. Thereafter, the warming airflow WA flows aft along the axial direction A between theinner surface 182 of theforward end 112 and thebase plate 158 of theouter dome 118. Some of the warming airflow WA flows into theinner surface openings 198 and convection heat transfer occurs. That is, the warming airflow WA flows into theinner surface openings 198 and exchanges heat with thesidewalls 199 of the openings, thus effectively warming theforward end 112 of theouter liner 108. After flowing into and out of theinner surface openings 198, the warming airflow WA flows aft along the axial direction A through the radial gap GR (FIG. 3 ) and into thecombustion chamber 114. Flowing the warming airflow WA into and out of theblind warming openings 190, which in this embodiment includes theouter surface openings 196 and theinner surface openings 198, theforward end 112 of theouter liner 108 is actively warmed. Accordingly, during transient bursts of the engine, the transient response rate of theforward end 112 is improved thereby reducing the bending stress and strain on the liner. - It will be appreciated that the
forward end 106 of the inner liner 102 (FIG. 2 ) may be formed in the same or substantially the same manner as theforward end 112 of theouter liner 108 as described above. Further, it will be appreciated that theforward end 106 of theinner liner 102 may be attached to theinner dome section 116 in the same or substantially the same manner that theforward end 112 of theouter liner 108 is attached to theouter dome section 118. -
FIG. 5 provides a close up, cross-sectional view of one exemplary embodiment ofcombustor assembly 100 depicting forward end 112 ofouter liner 108 attached toouter dome 118. Further,FIG. 5 depicts theforward end 112 ofouter liner 108 definingblind warming opening 190, which in the depicted embodiment ofFIG. 5 is a blind warming slit 200. Like the blind warming openings depicted in the embodiment ofFIG. 4 , the warming slit 200 defined by theforward end 112 increases the thermal response of theforward end 112 of theouter liner 108. Although only one (1) warming slit 200 is shown inFIG. 5 , in some embodiments theforward end 112 may define a plurality of warming slits 200. In particular, theforward end 112 of theouter liner 108 may define a plurality of warming slits 200 spaced apart from one another along the circumferential direction C. Moreover, although only one (1) slit is shown between theouter surface 178 and theinner surface 182 along the radial direction R, in some embodiments, theforward end 112 may define more than one warming slit 200 along the radial direction R. Theforward end 106 of theinner liner 102 can likewise define blind warming slits 200. The diameter of the blind warming slits 200 may range between or about between 0.020 and 0.10 inches (0.508-2.54 mm). The diameter of the mountingopenings 174 may range between or about between 0.400 and 0.800 inches (10.16-20.32 mm). As such, in such embodiments, the mountingopenings 174 have larger diameters than the blind warming slits 200. - As shown in
FIG. 5 , the warming slit 200 extends between anopening end 202 and aterminus end 204 along the axial direction A. For this embodiment, theradial interface surface 188 defines the openingend 202 of the warming slit 200. Moreover for this embodiment, the warming slit 200 extends along the axial direction A aft of the mountingopening 174. In this way, during operation of turbofan engine 10 (FIG. 1 ), theforward end 112 is warmed by warming airflow WA from theradial interface surface 188 to a point along the axial length of theforward end 112 that is aft of the mountingassembly 144 that couples theforward end 112 of theouter liner 108 with theouter dome 118. As the area of theforward end 112 that is axially aligned with and forward of the mountingassemblies 144 are most susceptible to thermal lag, flowing warming airflow WA into the warming slit 200 along this length reduces the bending stress and strain on the liner during transient operation of the turbofan engine 10. - In some exemplary embodiments, the warming slit 200 is defined by the
forward end 112 along the axial direction A from theradial interface surface 188 to a position that is aligned with or aft of anedge 143 along the axial direction A. Theedge 143, as shown inFIG. 5 , is a radial outermost point of a trailingedge 145 ofheat shield 142. As theheat shield 142 generally facilitates the flow ofcombustion gases 66 downstream of theforward end 112 of the outer liner 108 (FIG. 2 ), theforward end 112 is susceptible to thermal lag from itsradial interface surface 188 to a point along the axial direction A that is protected by theheat shield 142. Thus, by extending the warming slit 200 along this length effectively warms theforward end 112 during transient operation of the turbofan engine 10 (FIG. 1 ), and thus, bending stress and strain on the liner may be reduced. - In some exemplary embodiments, particularly where the
combustor assembly 100 does not includeheat shield 142 or similar structure, the warming slit 200 is defined by theforward end 112 along the axial direction A from theradial interface surface 188 to a position that is aligned with or aft of anedge 147 along the axial direction A. Theedge 147, as shown inFIG. 5 , is a radial outermost point of a trailingedge 149 ofouter dome 118. As theouter dome 118 generally shields theforward end 112 of theouter liner 108 fromcombustion gases 66, theforward end 112 is susceptible to thermal lag from itsradial interface surface 188 to a point along the axial direction A that is shielded by theouter dome 118. Thus, by extending the warming slit 200 along this length effectively warms theforward end 112 during transient operation of the turbofan engine 10 (FIG. 1 ), and thus, bending stress and strain on the liner may be reduced. - With reference to
FIG. 5 , warming airflow WA may flow into theblind warming openings 190, which in this embodiment is warming slit 200, as follows. As shown, during operation of turbofan engine 10 (FIG. 1 ), the warming airflow WA flows into theslot 122 defined by theouter dome 118. More particularly, the warming air WA flows between theyoke 160 and theouter surface 178 of theouter liner 108. In this example, the warming airflow WA is compressor discharge air (P3 air). After the warming airflow WA flows into theslot 122, the warming airflow WA flows around the most forward portion of theforward end 112 and flows into axial gap GA. As the warming airflow WA flows through the axial gap GA, some of the warming airflow WA flows into the openingend 202 of the warming slit 200. The warming airflow WA then travels forward along the axial direction A toward theterminus end 204 of the warming slit 200. As the warming airflow WA flows along the axial length of the warming slit 200, convection heat transfer occurs. That is, the warming airflow WA flowing through warming slit 200 exchanges heat with the walls 206 defining the warming slit 200, thus effectively warming theforward end 112 of theouter liner 108. After flowing into and out of the warming slit 200, the warming airflow WA flows aft along the axial direction A between theinner surface 182 of theouter liner 108 and thebase plate 158 and through the radial gap GR and into thecombustion chamber 114. Flowing the warming airflow WA into and out of theblind warming openings 190, which in this embodiment is warming slit 200 or slits, theforward end 112 of theouter liner 108 is actively warmed. Accordingly, during transient bursts of the engine, the transient response rate of theforward end 112 may be improved, which may reduce the bending stress and strain on the liner. - In some embodiments, forward end 112 of
outer liner 108 defines outer surface openings 196 (FIG. 4 ), inner surface openings 198 (FIG. 4 ), and warming slits 200 (FIG. 5 ). Moreover, it will be appreciated that theforward end 106 of theinner liner 102 can likewise include such configurations. -
FIG. 6 provides a close up, cross-sectional view of one exemplary embodiment ofcombustor assembly 100 depicting forward end 112 ofouter liner 108 attached toouter dome 118.FIG. 6 further depicts theouter dome 118 defining a plurality ofimpingement openings 210. In particular, theyolk 160 of theouter dome 118 defines theimpingement openings 210. As shown, theyolk 160 extends between anouter surface 161 and aninner surface 163 along the radial direction R, and for this embodiment, theimpingement openings 210 extend from theouter surface 161 to theinner surface 163 of theyolk 160. Thus, theimpingement openings 210 are through holes that extend through theyolk 160. - As further shown in
FIG. 6 , theimpingement openings 210 each extend between anouter end 212 and aninner end 214. The outer ends 212 of theimpingement openings 210 are defined by theouter surface 161 of theyolk 160 and the inner ends 214 of theimpingement openings 210 are defined by theinner surface 163 of theyolk 160. Theouter end 212 of eachimpingement opening 210 has an outer diameter DO and theinner end 214 of eachimpingement opening 210 has an inner diameter DI. For this embodiment, the outer diameters DO of theimpingement openings 210 are greater than the inner diameters DI. Stated differently, theimpingement openings 210 taper as they extend radially inward from theouter surface 161 to theinner surface 163 of theyolk 160. - During operation of the turbofan engine 10 (
FIG. 1 ), warming airflow WA (e.g., compressor discharge air) flows into theimpingement openings 210. In particular, the warming airflow WA enters theimpingement openings 210 through their respective outer ends 212. As the warming airflow WA flows inward along the radial direction R through theimpingement openings 210, the pressure of the warming airflow increases due to the converging geometry of theimpingement openings 210. Then, as the warming airflow WA exits the inner ends 214 of theimpingement openings 210, the flow rate of the warming airflow WA rapidly increases thereby creating a high velocity or jet stream of warming airflow WA. The streams of warming airflow WA flow inward along the radial direction R and impinges on theouter surface 178 of theforward end 112 of theouter liner 108. By impinging air onto theforward end 112, theforward end 112 is actively warmed and thus the transient thermal response rate of theforward end 112 is increased. The increased transient thermal response of theforward end 112 may reduce the bending stress and strain on the liner. - In some alternative embodiments, the inner diameters DI of the
impingement openings 210 are greater than the outer diameters DO. Stated alternatively, theimpingement openings 210 taper as they extend radially outward from theouter surface 161 to theinner surface 163 of theyolk 160. In such embodiments, the warming airflow WA flowing radially outward of theyolk 160 can more easily flow into theimpingement openings 210, among other benefits. In yet other alternative embodiments, the inner diameters DI of theimpingement openings 210 are the same as the outer diameters DO. That is, theimpingement openings 210 do not taper as they extend along the radial direction R. In further alternative embodiments, theimpingement openings 210 are angled with respect to the radial direction R. For instance, as one example, the inner ends 214 of theimpingement openings 210 may be positioned aft of their respective outer ends 212. In this way, the warming airflow WA may flow into theimpingement openings 210 more easily. -
FIG. 7 provides a close up, cross-sectional view of one exemplary embodiment ofcombustor assembly 100 depicting forward end 112 ofouter liner 108 attached toouter dome 118.FIG. 7 further depicts theouter dome 118 definingimpingement openings 210. Moreover,FIG. 7 also depicts animpingement jacket 220 attached to theouter dome 118. Theimpingement jacket 220 extends between aforward end 222 and an aft end (not shown) generally along the axial direction A. Generally, theimpingement jacket 220 spans the axial length of the outer liner 108 (i.e., from theforward end 112 to theaft end 110 of theouter liner 108; seeFIG. 2 ). As shown, theforward end 222 of theimpingement jacket 220 is attached to theyolk 160 of theouter dome 118. Theimpingement jacket 220 may be attached to theyolk 160 in any suitable manner. Theimpingement jacket 220 may extend annularly about the circumferential direction C or may extend along the circumferential direction C in segments. - As further shown in
FIG. 7 , like theouter dome 118, theimpingement jacket 220 defines a plurality ofimpingement openings 224. Theimpingement openings 224 may be configured in any suitable manner as noted above with respect to theimpingement openings 210 defined by theyolk 160 of theouter dome 118. The diameter of theblind warming openings 190 may range between or about between 0.020 and 0.080 inches (0.508-2.032 mm). Accordingly, during operation of turbofan engine 10 (FIG. 1 ), warming airflow WA may flow through theimpingement openings 224 defined by theimpingement jacket 220 and impinge on theouter surface 178 of theforward end 112 of theouter liner 108. -
FIG. 8 provides a close up, cross-sectional view of one exemplary embodiment ofcombustor assembly 100 depicting forward end 112 ofouter liner 108 attached toouter dome 118. Moreover,FIG. 8 depictsmetal grommet 172 of mountingassembly 144 positioned within mountingopening 174 and defining a plurality of warmingpassages 230. The warmingpassages 230 may be drilled into or otherwise machined into thegrommet 172. As shown, the warmingpassages 230 defined bygrommet 172 each extend between anouter end 232 and aninner end 234. For this embodiment, the warmingpassages 230 defined bygrommet 172 extend along the radial direction R. The outer ends 232 of the warmingpassages 230 are defined by theouter collar 176 of thegrommet 172, while the inner ends 234 of the warmingpassages 230 are defined by theinner collar 180 of thegrommet 172. Further, as the warmingpassages 230 extend through thegrommet 172, the warmingpassages 230 also extend through thebody 184 of thegrommet 172. Notably, as shown inFIG. 8 , the warmingpassages 230 extend through thebody 184 of thegrommet 172 such that the warmingpassages 230 extend along aradial face 113 of theforward end 112 of theouter liner 108. In this way, during operation of the turbofan engine 10 (FIG. 1 ), warming airflow WA may flow through the warmingpassages 230 and warm theradial face 113 of the forward 112, thereby actively warming theforward end 112. -
FIG. 9 provides a cross sectional view of thegrommet 172 ofFIG. 8 taken along line 9-9 ofFIG. 8 . As shown, thebushing 164 is generally cylindrical in shape and is positioned around theshank 168 of thepin 162. Thebushing 164 and pin 162 are positioned within and extend through a grommet opening 173 defined bygrommet 172. Thegrommet 172 is positioned in the mountingopening 174 defined by theforward end 112 of theouter liner 108. For this embodiment, as shown, the warmingpassages 230 are circumferentially spaced about theperimeter 185 of thebody 184 of thegrommet 172. Eachwarming passage 230 extends through theperimeter 185 of thebody 184 as a cutout and has a semi-circle shaped radial cross section and is, as noted above, positioned or defined bygrommet 172 such that the warmingpassages 230 are defined in part by thegrommet 172 and defined in part by theforward end 112 of theouter liner 108. This, as noted above, may allow for warming airflow WA to pass therethrough and actively warm theradial face 113 and theforward end 112 more generally. Thus, by definingwarming passages 230 ingrommet 172, the transient thermal response rate of theforward end 112 can be increased, and accordingly, the bending stress and strain on theouter liner 108 may be reduced. - In some alternative embodiments, additionally or alternatively to manufacturing the
grommet 172 with warmingpassages 230, thegrommet 172 may be sized such that a gap is defined between thegrommet 172 and theradial face 113 of theforward end 112. In this way, during operation of the turbofan engine 10 (FIG. 1), warming airflow WA may flow through the gap and warm theradial face 113, thereby actively warming theforward end 112 of theouter liner 108. -
FIG. 10 provides a close up, cross-sectional view of one exemplary embodiment ofcombustor assembly 100 depicting forward end 112 ofouter liner 108 attached toouter dome 118. Further,FIG. 10 depicts theforward end 112 having enhanced surfaces. In particular, for this embodiment, theouter surface 178 of theouter liner 108 has a heat transfer coefficient (HTC) enhancedsurface 240 and theinner surface 182 of theouter liner 108 likewise has an HTC enhancedsurface 240. By enhancing the surface of theouter surface 178 andinner surface 182 of theforward end 112 as shown inFIG. 10 , the surface area of theouter surface 178 is increased and the surface area of theinner surface 182 is likewise increased. Additionally, the enhancement of theouter surface 178 andinner surface 182 causes more turbulent flow over or across the surfaces. Thus, by enhancing the surfaces of theforward end 112, the transient thermal response of theforward end 112 may be increased. Additionally, as the enhanced surfaces do not affect the chargeable flow through the turbofan engine 10 (FIG. 1 ), there is no penalty on engine efficiency or performance. - As one example, the
enhanced surface 240 of theouter surface 178 and/or theinner surface 182 may be a rough bond coating. For instance, the coating may a suitable ceramic thermal and environmental barrier coating (TEBC) for CMC components. As another example, theenhanced surface 240 of theouter surface 178 and/or theinner surface 182 may have an undulating surface as shown inFIG. 10 . As yet another example, theenhanced surface 240 of theouter surface 178 and/or theinner surface 182 may be a grit blasted coating. For instance, the outer and/orinner surfaces forward end 112 may be blasted with a coarse grit. As a further example, theenhanced surface 240 of theouter surface 178 and/or theinner surface 182 may have a plurality of bumps having a height of about ten (10) millimeters. - Although the exemplary embodiments of the present disclosure were mainly discussed and illustrated using the outer liner and outer dome section of the combustor assembly, it will be appreciated that each exemplary aspect disclosed herein is applicable to the inner liner and inner dome section of the combustor assembly.
- 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)
Priority Applications (1)
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US15/860,835 US20190203940A1 (en) | 2018-01-03 | 2018-01-03 | Combustor Assembly for a Turbine Engine |
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Application Number | Priority Date | Filing Date | Title |
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US15/860,835 US20190203940A1 (en) | 2018-01-03 | 2018-01-03 | Combustor Assembly for a Turbine Engine |
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US20190203940A1 true US20190203940A1 (en) | 2019-07-04 |
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ID=67059418
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US15/860,835 Abandoned US20190203940A1 (en) | 2018-01-03 | 2018-01-03 | Combustor Assembly for a Turbine Engine |
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Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20220090788A1 (en) * | 2017-05-02 | 2022-03-24 | General Electric Company | Trapped vortex combustor for a gas turbine engine with a driver airflow channel |
US11859819B2 (en) | 2021-10-15 | 2024-01-02 | General Electric Company | Ceramic composite combustor dome and liners |
Citations (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5353587A (en) * | 1992-06-12 | 1994-10-11 | General Electric Company | Film cooling starter geometry for combustor lines |
US5653110A (en) * | 1991-07-22 | 1997-08-05 | General Electric Company | Film cooling of jet engine components |
US6098397A (en) * | 1998-06-08 | 2000-08-08 | Caterpillar Inc. | Combustor for a low-emissions gas turbine engine |
US6237344B1 (en) * | 1998-07-20 | 2001-05-29 | General Electric Company | Dimpled impingement baffle |
US6397603B1 (en) * | 2000-05-05 | 2002-06-04 | The United States Of America As Represented By The Secretary Of The Air Force | Conbustor having a ceramic matrix composite liner |
US20020152751A1 (en) * | 2001-04-19 | 2002-10-24 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor |
US20030177769A1 (en) * | 2002-03-21 | 2003-09-25 | Graves Charles B. | Counter swirl annular combustor |
US20040079082A1 (en) * | 2002-10-24 | 2004-04-29 | Bunker Ronald Scott | Combustor liner with inverted turbulators |
US20040118122A1 (en) * | 2002-12-20 | 2004-06-24 | Mitchell Krista Anne | Mounting assembly for the forward end of a ceramic matrix composite liner in a gas turbine engine combustor |
US20040123598A1 (en) * | 2002-12-31 | 2004-07-01 | General Electric Company | High temperature combustor wall for temperature reduction by optical reflection and process for manufacturing |
US20140157783A1 (en) * | 2012-12-10 | 2014-06-12 | General Electric Company | System for Protecting an Inner Wall of a Combustor |
US20140360196A1 (en) * | 2013-03-15 | 2014-12-11 | Rolls-Royce Corporation | Shell and tiled liner arrangement for a combustor |
US20170003026A1 (en) * | 2015-06-30 | 2017-01-05 | Rolls-Royce Corporation | Combustor tile |
-
2018
- 2018-01-03 US US15/860,835 patent/US20190203940A1/en not_active Abandoned
Patent Citations (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5653110A (en) * | 1991-07-22 | 1997-08-05 | General Electric Company | Film cooling of jet engine components |
US5353587A (en) * | 1992-06-12 | 1994-10-11 | General Electric Company | Film cooling starter geometry for combustor lines |
US6098397A (en) * | 1998-06-08 | 2000-08-08 | Caterpillar Inc. | Combustor for a low-emissions gas turbine engine |
US6237344B1 (en) * | 1998-07-20 | 2001-05-29 | General Electric Company | Dimpled impingement baffle |
US6397603B1 (en) * | 2000-05-05 | 2002-06-04 | The United States Of America As Represented By The Secretary Of The Air Force | Conbustor having a ceramic matrix composite liner |
US20020152751A1 (en) * | 2001-04-19 | 2002-10-24 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor |
US20030177769A1 (en) * | 2002-03-21 | 2003-09-25 | Graves Charles B. | Counter swirl annular combustor |
US20040079082A1 (en) * | 2002-10-24 | 2004-04-29 | Bunker Ronald Scott | Combustor liner with inverted turbulators |
US20040118122A1 (en) * | 2002-12-20 | 2004-06-24 | Mitchell Krista Anne | Mounting assembly for the forward end of a ceramic matrix composite liner in a gas turbine engine combustor |
US20040123598A1 (en) * | 2002-12-31 | 2004-07-01 | General Electric Company | High temperature combustor wall for temperature reduction by optical reflection and process for manufacturing |
US20140157783A1 (en) * | 2012-12-10 | 2014-06-12 | General Electric Company | System for Protecting an Inner Wall of a Combustor |
US20140360196A1 (en) * | 2013-03-15 | 2014-12-11 | Rolls-Royce Corporation | Shell and tiled liner arrangement for a combustor |
US20170003026A1 (en) * | 2015-06-30 | 2017-01-05 | Rolls-Royce Corporation | Combustor tile |
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20220090788A1 (en) * | 2017-05-02 | 2022-03-24 | General Electric Company | Trapped vortex combustor for a gas turbine engine with a driver airflow channel |
US11788725B2 (en) * | 2017-05-02 | 2023-10-17 | General Electric Company | Trapped vortex combustor for a gas turbine engine with a driver airflow channel |
US11859819B2 (en) | 2021-10-15 | 2024-01-02 | General Electric Company | Ceramic composite combustor dome and liners |
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