US20130081401A1 - Impingement cooling of combustor liners - Google Patents
Impingement cooling of combustor liners Download PDFInfo
- Publication number
- US20130081401A1 US20130081401A1 US13/250,274 US201113250274A US2013081401A1 US 20130081401 A1 US20130081401 A1 US 20130081401A1 US 201113250274 A US201113250274 A US 201113250274A US 2013081401 A1 US2013081401 A1 US 2013081401A1
- Authority
- US
- United States
- Prior art keywords
- nozzle
- turbine engine
- gas turbine
- liner
- nozzles
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Images
Classifications
-
- 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
- F23R3/06—Arrangement of apertures along the flame tube
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
-
- 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 disclosure relates generally to systems and methods of impingement cooling a combustor liner of a gas turbine engine.
- a type of combustor liner called a double-walled liner, includes an inner liner that encloses a volume where combustion occurs and an outer liner that surrounds the inner liner.
- An annular space between the inner liner and the outer liner assists in the cooling of the liner.
- film cooling air in the annular space is directed into the combustor through holes in the inner liner to mix with the hot combustion gases within. The air absorbs the heat from the inner liner as it flows therethrough.
- a gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine.
- the double-walled liner may extend from an upstream end to a downstream end.
- the gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end.
- the plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end.
- the at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.
- a method of impingement cooling a double-walled combustor liner of a gas turbine engine may extend from an upstream end to a downstream end and include an inner liner and an outer liner positioned radially outwards the inner liner.
- the method may include combusting a fuel in a combustor of the gas turbine engine, and directing cooling air through a plurality of nozzles that extend radially inwards through the outer liner to impinge upon and cool the inner liner.
- the cooling air may be directed such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end.
- the cooling air directed through at least one nozzle of the plurality of nozzles may exit the at least one nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.
- a gas turbine engine may include an impingement cooled double-walled liner.
- the double-walled liner may include an inner liner and an outer liner disposed around a combustion space of the turbine engine and extend from an upstream end to a downstream end.
- the gas turbine engine may also include a plurality of nozzles that extend radially inwards through the outer liner to direct cooling air towards the inner liner.
- Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end.
- Each nozzle of the plurality of nozzles may include multiple air holes arranged in a shower head pattern at the second end.
- FIG. 1 is an illustration of an exemplary disclosed gas turbine engine
- FIG. 2 is a cut-away illustration of an exemplary combustor system of the gas turbine engine of FIG. 1 ;
- FIG. 3 is a cross-sectional view of an embodiment of the outer combustor wall of the gas turbine engine of FIG. 1 ;
- FIG. 4A is a cross-sectional view of another embodiment of the outer combustor wall of the gas turbine engine of FIG. 1 ;
- FIG. 4B is a cross-sectional view of another embodiment of the outer combustor wall of the gas turbine engine of FIG. 1 ;
- FIG. 5A is a cross-sectional view of an embodiment of a nozzle of the gas turbine engine of FIG. 1 ;
- FIG. 5B is another cross-sectional view of the exemplary nozzle of FIG. 5A ;
- FIG. 6 is a cross-sectional view of another embodiment of a nozzle of the gas turbine engine of FIG. 1 ;
- FIG. 7 is a cross-sectional view of another embodiment of a nozzle of the gas turbine engine of FIG. 1 .
- FIG. 1 illustrates an exemplary gas turbine engine (GTE) 100 having a compressor system 10 , a combustor system 20 , a turbine system 70 , and an exhaust system 90 arranged lengthwise along an engine axis 98 .
- the compressor system 10 may compress air and deliver the compressed air to an enclosure 72 of the combustor system 20 .
- the compressed air may be mixed with a fuel and directed into a combustor 50 through one or more fuel injectors 30 .
- the fuel-air mixture may ignite and burn in the combustor 50 to produce combustion gases that may be directed to the turbine system 70 .
- the turbine system 70 may extract energy from these combustion gases, and direct the exhaust gases to the atmosphere through the exhaust system 90 .
- FIG. 2 is a cut-away view of combustor system 20 showing the combustor 50 .
- Combustor 50 includes an outer combustor wall 22 and an inner combustor wall 24 annularly disposed about the engine axis 98 .
- the outer and the inner combustor walls ( 22 , 24 ) are joined together at an upstream end by a dome assembly to define a combustion space 58 therebetween.
- the combustion space 58 is fluidly coupled to turbine system 70 at the downstream end.
- the plurality of fuel injectors 30 positioned on the dome assembly, direct the fuel-air mixture to the combustion space 58 for combustion.
- This fuel-air mixture burns in a combustion zone (proximate the upstream end) of the combustion space 58 to produce high pressure combustion gases that flow downstream towards the turbine system 70 .
- the combustion of fuel-air mixture within the combustion space 58 heats the outer and the inner combustor walls ( 22 , 24 ). For increased reliability and performance of GTE 100 , it is desirable to cool these walls.
- the outer combustor wall 22 includes an inner liner 22 b and an outer liner 22 a
- the inner combustor wall 24 includes an inner liner 24 b and an outer liner 24 a.
- the inner liners 22 b, 24 b are radially spaced apart from the outer liners 22 a, 24 a to define annular cooling spaces 26 , 28 between them. These cooling spaces 26 , 28 extend from an upstream end 44 to a downstream end 46 of the combustor 50 .
- the combustion in the combustion space 58 may create oscillations of pressure (pressure waves) within the combustion space 58 that causes radial expansion and contraction (bulging) of the inner liners 22 b, 24 b with respect to the outer liners 22 a, 24 a.
- the outer liners 22 a, 24 a include a plurality perforations 32 , 34 that direct high pressure air from the enclosure 72 to impinge on, and cool, the inner liners 22 b, 24 b.
- This technology of impingement cooling the combustor liners is referred to in the industry as Augmented Backside Cooled (ABC) technology. It is known that the use of ABC technology decreases the emission of pollutants into the atmosphere. It should be noted that the general configuration of combustor system 20 illustrated in FIG. 2 is exemplary only, and that several variations are possible.
- FIG. 3 is a cross-sectional schematic of the outer combustor wall 22 illustrating the impingement cooling of the inner liner 22 b.
- a high pressure stream of air (“air jets 36 ”) enters the cooling space 26 through perforations 32 on the outer liner 22 a. These air jets 36 impinge on, and cool, the inner liner 22 b. After impingement, the spent air stream flows towards the downstream end 46 to form the cross-flow air 38 that may be mixed with the combustion gases or discarded. It is known that cross-flow air 38 from the upstream end 44 , interacts with, and degrades the ability of the air jets 36 at the downstream end 46 to impinge on, and cool, the inner liner 22 b.
- the cross-flow air 38 from a first perforation 32 a may degrade the ability of the air jet 36 from a second perforation 32 b, downstream of the first perforation 32 a, to impinge on the region of the inner liner 22 b under the second perforation 32 b.
- the cross-flow air 38 from the first and second perforations 32 a, 32 b may collectively further degrade the cooling ability of an air jet 36 from a third perforation 32 c, further downstream of the first perforation 32 a, to cool the inner liner 22 b under the third perforation 32 c.
- some (or all) of the perforations 32 may include extended ports or nozzles 48 (see FIGS. 4A-7 ) to reduce the impact of the cross-flow air 38 from an upstream perforation 32 on a downstream air jet 36 .
- FIG. 4A illustrates a cross-sectional view of the outer combustor wall 22 illustrating nozzles 48 attached to the perforations 32 .
- the nozzles 48 may include air holes 62 that extend from a first end 66 , positioned in the enclosure 72 outside the outer liner 22 a, to a second end 68 positioned in the cooling space 26 inside the outer liner 22 a. These air holes 62 may direct the compressed air in the enclosure 72 (air jets 36 of FIG. 3 ) to impinge on the inner liner 22 b.
- the nozzles 48 may be a separate part attached to the outer liner 22 a (by any conventional attachment process, such as brazing, etc.) or may be a region of the outer liner 22 a that is bent towards the inner liner 22 b (such as for example, the rim of a perforation that is folded towards the inner liner 22 b ).
- the nozzles 48 may be arranged on the outer liner 22 a such that a radial gap (t) between the second end 68 of a nozzle 48 and the inner liner 22 b decreases from the upstream end 44 to the downstream end 46 (that is, t a >t b >t c >t d >t e ).
- the length of the nozzles 48 may progressively increase from the upstream end 44 to the downstream end 46 . Because the air jets 36 enter the cooling space 26 closer to the inner liner 22 b at the downstream end 46 , the effect of the cross-flow air 38 from the upstream air jets 36 on the downstream air jets 36 will be lower. In some embodiments, substantially all the rows of perforations on the outer liner 22 a will include nozzles 48 , while in other embodiments, only selected rows of perforations along the length of the outer liner 22 a will include nozzles 48 .
- the perforations 34 on the inner combustor wall 24 will also include nozzles 48 so that the air jets 36 enter the cooling space 28 closer to the inner liner 24 b at the downstream end 46 than at the upstream end 44 .
- FIG. 4A illustrates the radial gap (t) between the nozzles 48 and the inner liner 22 b as decreasing substantially linearly from the upstream end 44 to the downstream end 46 , this is only exemplary.
- the radial gap (t) may vary in any manner (such as, for example, decrease exponentially from the upstream end to the downstream end).
- the radial gaps (t) of selected adjacent nozzles 48 may be substantially the same (such as, for example, t a ⁇ t b >t c >t d ⁇ t e ).
- only perforations 32 in selected regions of the outer liner 22 a may include nozzles 48 to direct the air jets 36 in these regions closer to the inner liner 22 b.
- nozzles 48 may only be included in a few rows of perforations 32 at the downstream end 46 in applications where only the air jets 36 from those few rows are detrimentally affected by the cross-flow air 38 from the upstream end 44 .
- the radial gap (t) between these nozzles 48 and the inner liner 22 b may be substantially the same (that is, t a ⁇ t b ).
- the radial gap (t) may instead be decreased by decreasing the distance between the inner liner 22 b and the outer liner 22 a (that is, the thickness of the cooling space 26 ) from the upstream end 44 to the downstream end 46 .
- the pressure pulses generated in the combustion space 58 during combustion may cause portions of the inner liner 22 b to bulge outwards toward the nozzles 48 in corresponding portions of the outer liner 22 a.
- Contact between the inner liner 22 b and a nozzle 48 may restrict, or even block, air flow (air jets 36 ) through the nozzle 48 , and result in uneven cooling of the liner.
- Some embodiments of the nozzles 48 of the current disclosure may be configured to allow the air flow to continue even when they are in contact with the inner liner 22 b.
- FIGS. 5A and 5B are cross-sectional illustrations of an exemplary embodiment of a nozzle 48 A of the current disclosure.
- FIG. 5A illustrates a cross-sectional view along a plane parallel to a longitudinal axis 88 of nozzle 48 A
- FIG. 5B illustrates a cross-sectional view along a plane transverse to the longitudinal axis 88 .
- One or more air holes 62 may direct compressed air out of nozzle 48 A at second end 68 .
- the one or more air holes 62 may form a shower head pattern of air holes 62 at the second end 68 .
- all the air holes 62 may extend from the first end 66 to the second end 68 , while in other embodiments (as illustrated in FIG. 5A ), a single air hole 62 that extends from the first end 66 may be divided into multiple air holes 62 to form a shower head pattern at the second end 68 .
- the single air hole may be divided into multiple air holes anywhere along the length of nozzle 48 A.
- the single air hole may be divided into multiple air holes proximate the second end 68 .
- multiple (for example, 2, 3, 4, 5, 6 etc.) air holes 62 may be positioned symmetrically around the longitudinal axis 88 at the second end 68 .
- the bulging liner may contact a central portion (proximate longitudinal axis 88 ) of the second end 68 of a nozzle 48 . And, since the air holes 62 are distributed around the central portion, some or all of the air holes 62 may remain unblocked by the bulging inner liner 22 b, 24 b. Even if flow through some of the air holes 62 is restricted (or even blocked) by the contacting inner liner 22 b, 24 b, the flow though the remaining air holes 62 may provide sufficient cooling for the inner liner 22 b, 24 b. Thus, a shower head pattern of air holes 62 in nozzle 48 A may allow air flow to continue through at least some of the air holes 62 when there is contact between the nozzle 48 A and the inner liner 22 b , 24 b.
- nozzle 48 A may include one or more projections 74 that project outwards from the second end 68 . In some embodiments, at least one of these projections 74 may be located between the outlets of the multiple air holes 62 at the second end 68 . Other projections (if any) may be located anywhere on, or proximate, the second end 68 . For instance, in some embodiments, the projections 74 may be substantially evenly distributed on the second end 68 of the nozzle 48 A. These projections 74 may contact a bulging inner liner 22 b, 24 b and act as a standoff to allow air flow through the air holes 62 of the nozzle 48 A. The projections 74 may have any shape and size.
- arc-shaped projections may extend towards the inner liner 22 b, 24 b from the periphery of nozzle 48 A.
- the projections 74 may have a rounded edge to reduce bearing stresses on the inner liner 22 b, 24 b during contact.
- other features such as grooves, cut-outs, etc. that allow air from the air holes 62 to exit out of the nozzle 48 A when there is contact between the nozzle 48 A and the inner liner 22 b, 24 b, may be provided.
- the shape of a nozzle 48 A may include a projecting region on the second end 68 .
- the second end 68 of an exemplary nozzle 48 B may have a curved shape with a projecting central region.
- the projecting central region may act as the projection 74 that contacts a bulging inner liner 22 b, 24 b.
- additional projections 74 may also be provided on second end 68 of nozzle 48 B.
- central axes ( 64 a, 64 b, 64 c, 64 d, etc.) of the multiple air holes 62 may be substantially parallel to the longitudinal axis 88 .
- the central axis of an air hole 62 may be inclined with respect to the longitudinal axis 88 of the nozzle.
- FIG. 7 illustrates an embodiment of a nozzle 48 C in with the central axes 64 a, 64 b of the air holes 62 a and 62 b make angles ⁇ a and ⁇ b , respectively, with respect to the longitudinal axis 88 .
- angles ⁇ a and ⁇ b may have the same or different magnitudes.
- a bulging inner liner 22 b, 24 b may contact the central portion of the second end 68 and allow air flow through the inclined air holes 62 even in the absence of a projection at the second end 68 .
- one or more projections 74 may also be provided at the second end 68 of nozzle 48 C to act as a stand-off.
- the inclined air holes 62 may also allow the air flowing through them to diverge and impinge on a larger area of the inner liner 22 b, 24 b.
- nozzle such as, for example nozzles 48 , 48 A, 48 B, 48 C, etc.
- a nozzle having one air hole 62 such as nozzle 48 of FIGS. 4A and 4B
- a nozzle having multiple inclined air holes 62 such as nozzle 48 C of FIG. 7
- nozzle 48 C of FIG. 7 may be used where the possibility of contact with the inner liner 22 b, 24 b exists.
- contact between a nozzle 48 and the inner liner 22 b, 24 b is described as being a result of a pressure wave in the combustor 50 that causes a portion of the inner liner 22 b , 24 b to bulge and contact one or more nozzles 48 on the outer liner 22 a, 24 b, this is only exemplary. In some applications, vibration of the combustor 50 may cause contact between the inner liner 22 b, 24 b and the nozzles 48 . Contact between a nozzle 48 and the inner liner 22 b, 24 b can occur for various other reasons, and the disclosed system can be used to provide continuous air flow through the nozzles 48 during contact that occurs for any reason.
- the disclosed systems and methods of impingement cooling a cylinder liner may be applicable to any turbine engine to reliably and effectively cool the cylinder liner.
- the disclosed system of impingement cooling is configured to prevent the impingement air flow from being blocked as a result of dimensional changes of the combustor liner during operation of the turbine engine. The operation of a gas turbine engine using a disclosed system of impingement cooling will now be explained.
- air may be drawn into compressor section 10 and compressed. This compressed air may then be directed to enclosure 72 around the combustor 50 .
- the combustor may enclose a combustion space 58 bounded by a double-walled liner (including inner liners 22 b, 24 b and outer liners 22 a, 24 a ). A portion of the compressed air may be mixed with fuel and combusted in the combustion space 58 . The combustion heats the inner liners 22 b, 24 b of the combustor 50 .
- a portion of the compressed air in the enclosure 72 is directed though the perforations 32 , 34 on the outer liner 22 a, 24 a to impinge on, and cool, the hot inner liner 22 b, 24 b ( FIGS. 3 ).
- nozzles 48 are provided on some or all the perforations 32 , 34 . These nozzles 48 deliver the impingement air jets closer to the inner liner 22 b, 24 b at the downstream end 46 of the combustor 50 and thereby reduce the effect of the cross-flow air on the cooling effectiveness of the downstream air jets.
- the air jets may be provided in a shower head pattern at the tip of the nozzles 48 .
- a shower head pattern of air jets may allow some of the air jets to continue to impinge on, and cool, the inner liner 22 b, 24 b even when a bulging inner liner contacts and blocks some of the air jets.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine. The double-walled liner may extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. The plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end. The at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.
Description
- The present disclosure relates generally to systems and methods of impingement cooling a combustor liner of a gas turbine engine.
- Combustor liners of gas turbine engines are exposed to high temperatures of combustion and therefore require cooling. A type of combustor liner, called a double-walled liner, includes an inner liner that encloses a volume where combustion occurs and an outer liner that surrounds the inner liner. An annular space between the inner liner and the outer liner assists in the cooling of the liner. There are various methods that are employed to cool combustion liners during operation of the engine. These method include film cooling and jet impingement cooling. In film cooling, air in the annular space is directed into the combustor through holes in the inner liner to mix with the hot combustion gases within. The air absorbs the heat from the inner liner as it flows therethrough. In jet impingement cooling, air jets impinge upon and cool the back surface of the inner liner. These air jets may be directed to the back surface of the inner liner through an array of holes on the outer liner. After impinging on the back surface of the inner liner, the spent cooling air flows downstream through the annular space. This spent air flow, called cross-flow, is known to degrade the cooling ability of downstream air jets.
- U.S. Patent Application No. 2008/0271458 to Ekkad et al. (the '458 publication) describes an impingement cooled liner with ports extending from the outer liner to the inner liner to reduce the effects of cross-flow. While the extended ports of the '458 publication may reduce the effects of cross-flow, they may have limitations. For instance, dimensional changes during operation of the turbine engine may force portions of the inner liner against the extended ports preventing air flow therethrough. The systems and methods of the current disclosure are directed to overcoming one or more of the problems set forth above.
- In one aspect, a gas turbine engine is disclosed. The gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine. The double-walled liner may extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. The plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end. The at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.
- In another aspect, a method of impingement cooling a double-walled combustor liner of a gas turbine engine is disclosed. The double-walled liner may extend from an upstream end to a downstream end and include an inner liner and an outer liner positioned radially outwards the inner liner. The method may include combusting a fuel in a combustor of the gas turbine engine, and directing cooling air through a plurality of nozzles that extend radially inwards through the outer liner to impinge upon and cool the inner liner. The cooling air may be directed such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end. The cooling air directed through at least one nozzle of the plurality of nozzles may exit the at least one nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.
- In yet another aspect, a gas turbine engine is disclosed. The gas turbine engine may include an impingement cooled double-walled liner. The double-walled liner may include an inner liner and an outer liner disposed around a combustion space of the turbine engine and extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles that extend radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. Each nozzle of the plurality of nozzles may include multiple air holes arranged in a shower head pattern at the second end.
-
FIG. 1 is an illustration of an exemplary disclosed gas turbine engine; -
FIG. 2 is a cut-away illustration of an exemplary combustor system of the gas turbine engine ofFIG. 1 ; -
FIG. 3 is a cross-sectional view of an embodiment of the outer combustor wall of the gas turbine engine ofFIG. 1 ; -
FIG. 4A is a cross-sectional view of another embodiment of the outer combustor wall of the gas turbine engine ofFIG. 1 ; -
FIG. 4B is a cross-sectional view of another embodiment of the outer combustor wall of the gas turbine engine ofFIG. 1 ; -
FIG. 5A is a cross-sectional view of an embodiment of a nozzle of the gas turbine engine ofFIG. 1 ; -
FIG. 5B is another cross-sectional view of the exemplary nozzle ofFIG. 5A ; -
FIG. 6 is a cross-sectional view of another embodiment of a nozzle of the gas turbine engine ofFIG. 1 ; and -
FIG. 7 is a cross-sectional view of another embodiment of a nozzle of the gas turbine engine ofFIG. 1 . -
FIG. 1 illustrates an exemplary gas turbine engine (GTE) 100 having acompressor system 10, acombustor system 20, aturbine system 70, and anexhaust system 90 arranged lengthwise along anengine axis 98. Thecompressor system 10 may compress air and deliver the compressed air to anenclosure 72 of thecombustor system 20. The compressed air may be mixed with a fuel and directed into acombustor 50 through one ormore fuel injectors 30. The fuel-air mixture may ignite and burn in thecombustor 50 to produce combustion gases that may be directed to theturbine system 70. Theturbine system 70 may extract energy from these combustion gases, and direct the exhaust gases to the atmosphere through theexhaust system 90. -
FIG. 2 is a cut-away view ofcombustor system 20 showing thecombustor 50. Combustor 50 includes anouter combustor wall 22 and aninner combustor wall 24 annularly disposed about theengine axis 98. The outer and the inner combustor walls (22, 24) are joined together at an upstream end by a dome assembly to define acombustion space 58 therebetween. - The
combustion space 58 is fluidly coupled toturbine system 70 at the downstream end. The plurality offuel injectors 30, positioned on the dome assembly, direct the fuel-air mixture to thecombustion space 58 for combustion. This fuel-air mixture burns in a combustion zone (proximate the upstream end) of thecombustion space 58 to produce high pressure combustion gases that flow downstream towards theturbine system 70. The combustion of fuel-air mixture within thecombustion space 58 heats the outer and the inner combustor walls (22, 24). For increased reliability and performance ofGTE 100, it is desirable to cool these walls. Theouter combustor wall 22 includes aninner liner 22 b and anouter liner 22 a, and theinner combustor wall 24 includes aninner liner 24 b and anouter liner 24 a. Theinner liners outer liners annular cooling spaces spaces upstream end 44 to adownstream end 46 of thecombustor 50. The combustion in thecombustion space 58 may create oscillations of pressure (pressure waves) within thecombustion space 58 that causes radial expansion and contraction (bulging) of theinner liners outer liners outer liners enclosure 72 to impinge on, and cool, theinner liners combustor system 20 illustrated inFIG. 2 is exemplary only, and that several variations are possible. -
FIG. 3 is a cross-sectional schematic of theouter combustor wall 22 illustrating the impingement cooling of theinner liner 22 b. A high pressure stream of air (“air jets 36”) enters the coolingspace 26 throughperforations 32 on theouter liner 22 a. Theseair jets 36 impinge on, and cool, theinner liner 22 b. After impingement, the spent air stream flows towards thedownstream end 46 to form thecross-flow air 38 that may be mixed with the combustion gases or discarded. It is known thatcross-flow air 38 from theupstream end 44, interacts with, and degrades the ability of theair jets 36 at thedownstream end 46 to impinge on, and cool, theinner liner 22 b. For instance, thecross-flow air 38 from afirst perforation 32 a may degrade the ability of theair jet 36 from asecond perforation 32 b, downstream of thefirst perforation 32 a, to impinge on the region of theinner liner 22 b under thesecond perforation 32 b. Similarly, thecross-flow air 38 from the first andsecond perforations air jet 36 from athird perforation 32 c, further downstream of thefirst perforation 32 a, to cool theinner liner 22 b under thethird perforation 32 c. In some embodiments, some (or all) of theperforations 32 may include extended ports or nozzles 48 (seeFIGS. 4A-7 ) to reduce the impact of thecross-flow air 38 from anupstream perforation 32 on adownstream air jet 36. -
FIG. 4A illustrates a cross-sectional view of theouter combustor wall 22 illustratingnozzles 48 attached to theperforations 32. Thenozzles 48 may includeair holes 62 that extend from afirst end 66, positioned in theenclosure 72 outside theouter liner 22 a, to asecond end 68 positioned in the coolingspace 26 inside theouter liner 22 a. These air holes 62 may direct the compressed air in the enclosure 72 (air jets 36 ofFIG. 3 ) to impinge on theinner liner 22 b. Thenozzles 48 may be a separate part attached to theouter liner 22 a (by any conventional attachment process, such as brazing, etc.) or may be a region of theouter liner 22 a that is bent towards theinner liner 22 b (such as for example, the rim of a perforation that is folded towards theinner liner 22 b). Thenozzles 48 may be arranged on theouter liner 22 a such that a radial gap (t) between thesecond end 68 of anozzle 48 and theinner liner 22 b decreases from theupstream end 44 to the downstream end 46 (that is, ta>tb>tc>td>te). To achieve this decreasing radial gap (t) from theupstream end 44 to thedownstream end 46, in some embodiments (as illustrated inFIG. 4A ), the length of thenozzles 48 may progressively increase from theupstream end 44 to thedownstream end 46. Because theair jets 36 enter thecooling space 26 closer to theinner liner 22 b at thedownstream end 46, the effect of thecross-flow air 38 from theupstream air jets 36 on thedownstream air jets 36 will be lower. In some embodiments, substantially all the rows of perforations on theouter liner 22 a will includenozzles 48, while in other embodiments, only selected rows of perforations along the length of theouter liner 22 a will includenozzles 48. In some embodiments, theperforations 34 on the inner combustor wall 24 (seeFIG. 2 ) will also includenozzles 48 so that theair jets 36 enter thecooling space 28 closer to theinner liner 24 b at thedownstream end 46 than at theupstream end 44. - Although
FIG. 4A illustrates the radial gap (t) between thenozzles 48 and theinner liner 22 b as decreasing substantially linearly from theupstream end 44 to thedownstream end 46, this is only exemplary. In general, the radial gap (t) may vary in any manner (such as, for example, decrease exponentially from the upstream end to the downstream end). In some embodiments, although the radial gap (t) may generally decrease from theupstream end 44 to thedownstream end 46, the radial gaps (t) of selectedadjacent nozzles 48 may be substantially the same (such as, for example, ta≈tb>tc>td≈te). - In some embodiments, as illustrated in
FIG. 4B , only perforations 32 in selected regions of theouter liner 22 a may includenozzles 48 to direct theair jets 36 in these regions closer to theinner liner 22 b. For example, in some embodiments,nozzles 48 may only be included in a few rows ofperforations 32 at thedownstream end 46 in applications where only theair jets 36 from those few rows are detrimentally affected by thecross-flow air 38 from theupstream end 44. In some embodiments (as illustrated inFIG. 4B ), the radial gap (t) between thesenozzles 48 and theinner liner 22 b may be substantially the same (that is, ta≈tb). AlthoughFIGS. 4A and 4B illustrate embodiments, wherenozzles 48 are used to decrease the radial gap (t) from theupstream end 44 to thedownstream end 46, it is contemplated that in some embodiments, the radial gap (t) may instead be decreased by decreasing the distance between theinner liner 22 b and theouter liner 22 a (that is, the thickness of the cooling space 26) from theupstream end 44 to thedownstream end 46. - In some applications, the pressure pulses generated in the
combustion space 58 during combustion may cause portions of theinner liner 22 b to bulge outwards toward thenozzles 48 in corresponding portions of theouter liner 22 a. Contact between theinner liner 22 b and anozzle 48 may restrict, or even block, air flow (air jets 36) through thenozzle 48, and result in uneven cooling of the liner. Some embodiments of thenozzles 48 of the current disclosure may be configured to allow the air flow to continue even when they are in contact with theinner liner 22 b. -
FIGS. 5A and 5B are cross-sectional illustrations of an exemplary embodiment of anozzle 48A of the current disclosure.FIG. 5A illustrates a cross-sectional view along a plane parallel to alongitudinal axis 88 ofnozzle 48A, andFIG. 5B illustrates a cross-sectional view along a plane transverse to thelongitudinal axis 88. In the discussion that follows, reference will be made to bothFIGS. 5A and 5B . One ormore air holes 62 may direct compressed air out ofnozzle 48A atsecond end 68. In some embodiments, as illustrated inFIG. 5A , the one ormore air holes 62 may form a shower head pattern ofair holes 62 at thesecond end 68. In some embodiments, all the air holes 62 may extend from thefirst end 66 to thesecond end 68, while in other embodiments (as illustrated inFIG. 5A ), asingle air hole 62 that extends from thefirst end 66 may be divided intomultiple air holes 62 to form a shower head pattern at thesecond end 68. The single air hole may be divided into multiple air holes anywhere along the length ofnozzle 48A. In some embodiments, as illustrated inFIG. 5A , the single air hole may be divided into multiple air holes proximate thesecond end 68. In some embodiments, multiple (for example, 2, 3, 4, 5, 6 etc.) air holes 62 may be positioned symmetrically around thelongitudinal axis 88 at thesecond end 68. If theinner liner second end 68 of anozzle 48. And, since the air holes 62 are distributed around the central portion, some or all of the air holes 62 may remain unblocked by the bulginginner liner inner liner inner liner air holes 62 innozzle 48A may allow air flow to continue through at least some of the air holes 62 when there is contact between thenozzle 48A and theinner liner - In some embodiments,
nozzle 48A may include one ormore projections 74 that project outwards from thesecond end 68. In some embodiments, at least one of theseprojections 74 may be located between the outlets of themultiple air holes 62 at thesecond end 68. Other projections (if any) may be located anywhere on, or proximate, thesecond end 68. For instance, in some embodiments, theprojections 74 may be substantially evenly distributed on thesecond end 68 of thenozzle 48A. Theseprojections 74 may contact a bulginginner liner nozzle 48A. Theprojections 74 may have any shape and size. For instance, in some embodiments, arc-shaped projections may extend towards theinner liner nozzle 48A. In some embodiments, as illustrated inFIG. 5A , theprojections 74 may have a rounded edge to reduce bearing stresses on theinner liner nozzle 48A when there is contact between thenozzle 48A and theinner liner discrete projections 74 on thesecond end 68, in some embodiments, the shape of anozzle 48A may include a projecting region on thesecond end 68. For example, as illustrated inFIG. 6 , thesecond end 68 of anexemplary nozzle 48B may have a curved shape with a projecting central region. In these embodiments, the projecting central region may act as theprojection 74 that contacts a bulginginner liner additional projections 74 may also be provided onsecond end 68 ofnozzle 48B. - In some embodiments (as illustrated in
nozzles FIGS. 5A-6 ), central axes (64 a, 64 b, 64 c, 64 d, etc.) of themultiple air holes 62 may be substantially parallel to thelongitudinal axis 88. However, in other embodiments, the central axis of anair hole 62 may be inclined with respect to thelongitudinal axis 88 of the nozzle.FIG. 7 illustrates an embodiment of anozzle 48C in with thecentral axes longitudinal axis 88. The angles θa and θb may have the same or different magnitudes. In these embodiments, a bulginginner liner second end 68 and allow air flow through the inclined air holes 62 even in the absence of a projection at thesecond end 68. In some embodiments, one ormore projections 74 may also be provided at thesecond end 68 ofnozzle 48C to act as a stand-off. The inclined air holes 62 may also allow the air flowing through them to diverge and impinge on a larger area of theinner liner - Any type of nozzle (such as, for
example nozzles nozzle 48 ofFIGS. 4A and 4B ) may be used in areas where the possibility of contact with theinner liner nozzle 48C ofFIG. 7 ) may be used where the possibility of contact with theinner liner nozzle 48 and theinner liner combustor 50 that causes a portion of theinner liner more nozzles 48 on theouter liner combustor 50 may cause contact between theinner liner nozzles 48. Contact between anozzle 48 and theinner liner nozzles 48 during contact that occurs for any reason. - The disclosed systems and methods of impingement cooling a cylinder liner may be applicable to any turbine engine to reliably and effectively cool the cylinder liner. The disclosed system of impingement cooling is configured to prevent the impingement air flow from being blocked as a result of dimensional changes of the combustor liner during operation of the turbine engine. The operation of a gas turbine engine using a disclosed system of impingement cooling will now be explained.
- With reference to
FIGS. 1 and 2 , during operation ofGTE 100, air may be drawn intocompressor section 10 and compressed. This compressed air may then be directed toenclosure 72 around thecombustor 50. The combustor may enclose acombustion space 58 bounded by a double-walled liner (includinginner liners outer liners combustion space 58. The combustion heats theinner liners combustor 50. A portion of the compressed air in theenclosure 72 is directed though theperforations outer liner inner liner FIGS. 3 ). To reduce the impact ofcross-flow air 38, from upstream perforations, on the cooling effectiveness of downstream perforations,nozzles 48 are provided on some or all theperforations nozzles 48 deliver the impingement air jets closer to theinner liner downstream end 46 of thecombustor 50 and thereby reduce the effect of the cross-flow air on the cooling effectiveness of the downstream air jets. To reduce the possibility of the air jets being blocked by dimensional variations of the liner walls during operation of the turbine engine (such as bulging of theinner liner nozzles 48. A shower head pattern of air jets may allow some of the air jets to continue to impinge on, and cool, theinner liner - It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed impingement cooling system and method. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed cooling system. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.
Claims (20)
1. A gas turbine engine, comprising:
an impingement cooled double-walled liner, including an inner liner and an outer liner, disposed around a combustion space of the turbine engine and extending from an upstream end to a downstream end; and
a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner, each nozzle of the plurality of nozzles extending radially inwards from a first distal end to a second proximal end, the plurality of nozzles being arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end,
wherein, at least one nozzle of the plurality of nozzles includes multiple air holes at the second end.
2. The gas turbine engine of claim 1 , wherein the at least one nozzle further includes a longitudinal axis extending from the first end to the second end and each air hole of the multiple air holes includes a central axis, the multiple air holes being symmetrically arranged about the longitudinal axis.
3. The gas turbine engine of claim 2 , wherein the central axis of each air hole is substantially parallel to the longitudinal axis.
4. The gas turbine engine of claim 2 , wherein the central axis of each air hole is inclined with respect to the longitudinal axis such that the cooling air exiting the at least one nozzle diverges.
5. The gas turbine engine of claim 1 , wherein the multiple air holes in the at least nozzle is arranged in a shower head pattern at the second end.
6. The gas turbine engine of claim 1 , wherein the second end of the at least one nozzle includes a projection that extends towards the inner liner.
7. The gas turbine engine of claim 6 , wherein the projection is centrally positioned on the second end and each air hole of the multiple air holes is symmetrically positioned about the projection.
8. The gas turbine engine of claim 1 , wherein the second end of the at least one nozzle is curved such that a central portion of the second end forms a proximal-most portion of the nozzle.
9. The gas turbine engine of claim 1 , wherein the radial gap decreases substantially linearly from the upstream end to the downstream end.
10. A method of impingement cooling a double-walled combustor liner of a gas turbine engine, the double-walled liner extending from an upstream end to a downstream end and including an inner liner and an outer liner positioned radially outwards the inner liner, comprising:
combusting a fuel in a combustor of the gas turbine engine; and
directing cooling air through a plurality of nozzles extending radially inwards through the outer liner to impinge upon and cool the inner liner, such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end, wherein the cooling air directed through at least one nozzle of the plurality of nozzles exit the at least nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.
11. The method of claim 10 , wherein directing the cooling air includes directing the cooling air through the multiple air flow paths of the at least one nozzle such that the cooling air diverges.
12. The method of claim 10 , wherein directing the cooling air includes directing the cooling air though the multiple air flow paths of the at least nozzle such that the cooling air through each of the multiple air flow paths flow substantially parallel to one another.
13. A gas turbine engine, comprising:
an impingement cooled double-walled liner, including an inner liner and an outer liner, disposed around a combustion space of the turbine engine and extending from an upstream end to a downstream end; and
a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner, each nozzle of the plurality of nozzles extending radially inwards from a first distal end to a second proximal end, wherein each nozzle of the plurality of nozzles include multiple air holes arranged in a shower head pattern at the second end.
14. The gas turbine engine of claim 13 , wherein the plurality of nozzles are arranged such that a radial gap of the second end of a nozzle to the inner liner decreases as a function of distance from the upstream end to the downstream end.
15. The gas turbine engine of claim 13 , wherein the multiple air holes are symmetrically positioned about a longitudinal axis of each nozzle.
16. The gas turbine engine of claim 15 , wherein each air hole of the multiple air holes are inclined with respect to the longitudinal axis such that the cooling air exiting each nozzle diverges.
17. The gas turbine engine of claim 16 , wherein an inclination of each air hole of the multiple air holes with respect to the longitudinal axis is substantially the same.
18. The gas turbine engine of claim 13 , wherein the second end of each nozzle includes a projection that extends towards the inner liner.
19. The gas turbine engine of claim 18 , wherein the multiple air holes are symmetrically arranged about the projection.
20. The gas turbine engine of claim 13 , wherein the second end of each nozzle is curved such that a central portion of the second end forms a proximal-most portion of the nozzle.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/250,274 US20130081401A1 (en) | 2011-09-30 | 2011-09-30 | Impingement cooling of combustor liners |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/250,274 US20130081401A1 (en) | 2011-09-30 | 2011-09-30 | Impingement cooling of combustor liners |
Publications (1)
Publication Number | Publication Date |
---|---|
US20130081401A1 true US20130081401A1 (en) | 2013-04-04 |
Family
ID=47991348
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US13/250,274 Abandoned US20130081401A1 (en) | 2011-09-30 | 2011-09-30 | Impingement cooling of combustor liners |
Country Status (1)
Country | Link |
---|---|
US (1) | US20130081401A1 (en) |
Cited By (19)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20150361889A1 (en) * | 2014-06-11 | 2015-12-17 | Alstom Technology Ltd | Impingement cooled wall arrangement |
EP3098386A1 (en) * | 2015-05-29 | 2016-11-30 | General Electric Company | Impingement insert |
EP3133242A1 (en) * | 2015-08-17 | 2017-02-22 | General Electric Company | Manifold with impingement plate for thermal adjustment of a turbine component |
US9849510B2 (en) | 2015-04-16 | 2017-12-26 | General Electric Company | Article and method of forming an article |
US9957816B2 (en) | 2014-05-29 | 2018-05-01 | General Electric Company | Angled impingement insert |
US10087776B2 (en) | 2015-09-08 | 2018-10-02 | General Electric Company | Article and method of forming an article |
US20190099810A1 (en) * | 2013-03-15 | 2019-04-04 | United Technologies Corporation | Additive manufacturing baffles, covers, and dies |
US10253986B2 (en) | 2015-09-08 | 2019-04-09 | General Electric Company | Article and method of forming an article |
EP3477063A1 (en) * | 2017-10-26 | 2019-05-01 | MAN Energy Solutions SE | Flow engine with a specific impingement cooling assembly |
US20190162080A1 (en) * | 2017-11-30 | 2019-05-30 | United Technologies Corporation | Apparatus and method for mitigating particulate accumulation on a component of a gas turbine engine |
US10309228B2 (en) * | 2016-06-09 | 2019-06-04 | General Electric Company | Impingement insert for a gas turbine engine |
US10422235B2 (en) | 2014-05-29 | 2019-09-24 | General Electric Company | Angled impingement inserts with cooling features |
US10690055B2 (en) | 2014-05-29 | 2020-06-23 | General Electric Company | Engine components with impingement cooling features |
US10739087B2 (en) | 2015-09-08 | 2020-08-11 | General Electric Company | Article, component, and method of forming an article |
CN114857618A (en) * | 2021-02-03 | 2022-08-05 | 通用电气公司 | Combustor for a gas turbine engine |
US11525401B2 (en) | 2021-01-11 | 2022-12-13 | Honeywell International Inc. | Impingement baffle for gas turbine engine |
CN115507387A (en) * | 2021-06-07 | 2022-12-23 | 通用电气公司 | Combustor for a gas turbine engine |
CN115507384A (en) * | 2021-06-07 | 2022-12-23 | 通用电气公司 | Combustor for a gas turbine engine |
US11572801B2 (en) * | 2019-09-12 | 2023-02-07 | General Electric Company | Turbine engine component with baffle |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4916906A (en) * | 1988-03-25 | 1990-04-17 | General Electric Company | Breach-cooled structure |
US5586866A (en) * | 1994-08-26 | 1996-12-24 | Abb Management Ag | Baffle-cooled wall part |
US6000908A (en) * | 1996-11-05 | 1999-12-14 | General Electric Company | Cooling for double-wall structures |
US6237344B1 (en) * | 1998-07-20 | 2001-05-29 | General Electric Company | Dimpled impingement baffle |
US6484505B1 (en) * | 2000-02-25 | 2002-11-26 | General Electric Company | Combustor liner cooling thimbles and related method |
US20050097890A1 (en) * | 2003-08-29 | 2005-05-12 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor |
US20080271458A1 (en) * | 2007-03-01 | 2008-11-06 | Srinath Varadarajan Ekkad | Zero-Cross-Flow Impingement Via An Array of Differing Length, Extended Ports |
-
2011
- 2011-09-30 US US13/250,274 patent/US20130081401A1/en not_active Abandoned
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4916906A (en) * | 1988-03-25 | 1990-04-17 | General Electric Company | Breach-cooled structure |
US5586866A (en) * | 1994-08-26 | 1996-12-24 | Abb Management Ag | Baffle-cooled wall part |
US6000908A (en) * | 1996-11-05 | 1999-12-14 | General Electric Company | Cooling for double-wall structures |
US6237344B1 (en) * | 1998-07-20 | 2001-05-29 | General Electric Company | Dimpled impingement baffle |
US6484505B1 (en) * | 2000-02-25 | 2002-11-26 | General Electric Company | Combustor liner cooling thimbles and related method |
US20050097890A1 (en) * | 2003-08-29 | 2005-05-12 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor |
US20080271458A1 (en) * | 2007-03-01 | 2008-11-06 | Srinath Varadarajan Ekkad | Zero-Cross-Flow Impingement Via An Array of Differing Length, Extended Ports |
Cited By (26)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20190099810A1 (en) * | 2013-03-15 | 2019-04-04 | United Technologies Corporation | Additive manufacturing baffles, covers, and dies |
US9957816B2 (en) | 2014-05-29 | 2018-05-01 | General Electric Company | Angled impingement insert |
US10422235B2 (en) | 2014-05-29 | 2019-09-24 | General Electric Company | Angled impingement inserts with cooling features |
US10690055B2 (en) | 2014-05-29 | 2020-06-23 | General Electric Company | Engine components with impingement cooling features |
US10060352B2 (en) * | 2014-06-11 | 2018-08-28 | Ansaldo Energia Switzerland AG | Impingement cooled wall arrangement |
US20150361889A1 (en) * | 2014-06-11 | 2015-12-17 | Alstom Technology Ltd | Impingement cooled wall arrangement |
US9849510B2 (en) | 2015-04-16 | 2017-12-26 | General Electric Company | Article and method of forming an article |
JP2016223447A (en) * | 2015-05-29 | 2016-12-28 | ゼネラル・エレクトリック・カンパニイ | Article, component, and method of forming article |
EP3098386A1 (en) * | 2015-05-29 | 2016-11-30 | General Electric Company | Impingement insert |
US9976441B2 (en) | 2015-05-29 | 2018-05-22 | General Electric Company | Article, component, and method of forming an article |
US20180187557A1 (en) * | 2015-08-17 | 2018-07-05 | General Electric Company | Article and manifold for thermal adjustment of a turbine component |
EP3133242A1 (en) * | 2015-08-17 | 2017-02-22 | General Electric Company | Manifold with impingement plate for thermal adjustment of a turbine component |
US9995151B2 (en) | 2015-08-17 | 2018-06-12 | General Electric Company | Article and manifold for thermal adjustment of a turbine component |
JP2017040259A (en) * | 2015-08-17 | 2017-02-23 | ゼネラル・エレクトリック・カンパニイ | Article and manifold for thermal adjustment of turbine component |
US10739087B2 (en) | 2015-09-08 | 2020-08-11 | General Electric Company | Article, component, and method of forming an article |
US10087776B2 (en) | 2015-09-08 | 2018-10-02 | General Electric Company | Article and method of forming an article |
US10253986B2 (en) | 2015-09-08 | 2019-04-09 | General Electric Company | Article and method of forming an article |
US10309228B2 (en) * | 2016-06-09 | 2019-06-04 | General Electric Company | Impingement insert for a gas turbine engine |
EP3477063A1 (en) * | 2017-10-26 | 2019-05-01 | MAN Energy Solutions SE | Flow engine with a specific impingement cooling assembly |
US20190162080A1 (en) * | 2017-11-30 | 2019-05-30 | United Technologies Corporation | Apparatus and method for mitigating particulate accumulation on a component of a gas turbine engine |
US10995635B2 (en) * | 2017-11-30 | 2021-05-04 | Raytheon Technologies Corporation | Apparatus and method for mitigating particulate accumulation on a component of a gas turbine engine |
US11572801B2 (en) * | 2019-09-12 | 2023-02-07 | General Electric Company | Turbine engine component with baffle |
US11525401B2 (en) | 2021-01-11 | 2022-12-13 | Honeywell International Inc. | Impingement baffle for gas turbine engine |
CN114857618A (en) * | 2021-02-03 | 2022-08-05 | 通用电气公司 | Combustor for a gas turbine engine |
CN115507387A (en) * | 2021-06-07 | 2022-12-23 | 通用电气公司 | Combustor for a gas turbine engine |
CN115507384A (en) * | 2021-06-07 | 2022-12-23 | 通用电气公司 | Combustor for a gas turbine engine |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US20130081401A1 (en) | Impingement cooling of combustor liners | |
JP4124585B2 (en) | Combustor liner with selectively inclined cooling holes. | |
JP4433529B2 (en) | Multi-hole membrane cooled combustor liner | |
US10094564B2 (en) | Combustor dilution hole cooling system | |
US8544277B2 (en) | Turbulated aft-end liner assembly and cooling method | |
US10378774B2 (en) | Annular combustor with scoop ring for gas turbine engine | |
US9625152B2 (en) | Combustor heat shield for a gas turbine engine | |
US6470685B2 (en) | Combustion apparatus | |
EP2378200B1 (en) | Combustor liner cooling at transition duct interface and related method | |
US20110185739A1 (en) | Gas turbine combustors with dual walled liners | |
US9982890B2 (en) | Combustor dome heat shield | |
US20080115499A1 (en) | Combustor heat shield with variable cooling | |
US20080115498A1 (en) | Combustor liner and heat shield assembly | |
US20090120093A1 (en) | Turbulated aft-end liner assembly and cooling method | |
JP3924136B2 (en) | Gas turbine combustor | |
US8695351B2 (en) | Hula seal with preferential cooling having spring fingers and/or adjacent slots with different widths | |
US10753283B2 (en) | Combustor heat shield cooling hole arrangement | |
US9134028B2 (en) | Combustor for gas turbine engine | |
JP2004003835A (en) | Multihole patch for combustor liner of gas turbine engine | |
US9188336B2 (en) | Assemblies and apparatus related to combustor cooling in turbine engines | |
US9933161B1 (en) | Combustor dome heat shield | |
EP2230456A2 (en) | Combustion liner with mixing hole stub | |
US11204169B2 (en) | Combustor of gas turbine engine and method | |
CA2936200C (en) | Combustor cooling system | |
US10697636B2 (en) | Cooling a combustor heat shield proximate a quench aperture |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: SOLAR TURBINES INCORPORATED, CALIFORNIA Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:KIM, YONG WEON, MR.;REEL/FRAME:027001/0164 Effective date: 20110929 |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |