US20130000309A1 - System and method for adaptive impingement cooling - Google Patents
System and method for adaptive impingement cooling Download PDFInfo
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- US20130000309A1 US20130000309A1 US13/174,166 US201113174166A US2013000309A1 US 20130000309 A1 US20130000309 A1 US 20130000309A1 US 201113174166 A US201113174166 A US 201113174166A US 2013000309 A1 US2013000309 A1 US 2013000309A1
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- liner
- mounting support
- support
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- cooling air
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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
- F23R3/002—Wall structures
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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
- F23R3/02—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
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03042—Film cooled combustion chamber walls or domes
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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 invention relates to cooling systems, and in particular, to a system and method for adaptive impingement cooling for use in hot environments such as those found in gas turbine engines.
- Gas turbine engines operate according to a continuous Brayton cycle where a pressurized air and fuel mixture is ignited in a combustor to produce a flowing stream of hot gas. The air is compressed, used for combustion, expands through a turbine, and finally exits the engine. Some gas turbine engines also include an augmentation system downstream of the turbine, where fuel is also introduced and ignited to increase thrust. Most often, the temperature of the primary air is higher than the melting temperatures of the materials that form the combustor, turbine, and augmentation system components. As a result, adequate cooling is integral to the function of gas turbine engines.
- An adaptive cooling structure comprises a mounting support, a liner, and a spacer.
- the mounting support has a coolant aperture for directing cooling air through the support.
- the liner has a first surface facing away from the mounting support and a second surface facing towards the support.
- the liner is coupled to the mounting support, and the spacer is positioned between the support and the liner. The positioning of the spacer creates a chamber between the mounting support and the liner, thus allowing the cooling air to impinge on the second surface of the liner.
- the liner wall is configured to deflect away from the mounting support to expand the chamber, thus allowing the cooling air to further impinge on the second surface of the liner.
- FIG. 1 is a simplified cross-sectional view of an embodiment of a gas turbine engine which employs the adaptive impingement cooling system and method of the present invention.
- FIG. 2 is a partial isometric view of an embodiment of the adaptive cooling structure of the present invention.
- FIG. 3 is a cross-sectional view of the embodiment of the adaptive cooling structure in FIG. 2 at a non-hot spot location.
- FIG. 4 is a cross-sectional view of the embodiment of the adaptive cooling structure in FIG. 2 at a hot spot location.
- FIG. 5 is a graph showing preferred ranges of impingement effectiveness for designing the adaptive cooling structure of the present invention.
- FIG. 1 is a simplified cross-sectional view of mixed flow turbofan engine 10 which can employ the adaptive impingement cooling system and method of the present invention.
- Turbofan engine 10 includes augmentation system 12 , fan duct 14 , drive fan 16 , low pressure compressor 18 , high pressure compressor 20 , combustor 22 , high pressure turbine 24 , low pressure turbine 26 , and exhaust nozzle 28 .
- Drive fan 16 and low pressure compressor 18 are driven by low pressure turbine 26 with shaft 30 .
- High pressure compressor 20 is driven by high pressure turbine 24 with shaft 32 .
- High pressure compressor 20 , combustor 22 , high pressure turbine 24 and shaft 32 comprise the core of turbofan engine 10 .
- Augmentation system 12 includes augmenter duct 34 and augmenter liner 36 .
- Drive fan 16 is rotated by low pressure turbine 26 to accelerate A Ambient thereby producing a major portion of the thrust output of turbofan engine 10 .
- Accelerated A Ambient is divided into two streams of air: primary air A P and secondary air A S .
- Secondary air A S also known as bypass air, passes into fan duct 14 where it passes on to augmentation system 12 .
- Primary air A P also known as hot air, is a stream of air that is directed first into low pressure compressor 18 and then into high pressure compressor 20 .
- Pressurized primary air A P is then passed into combustor 22 where it is mixed with a fuel supply and ignited to produce the high energy gases used to turn high pressure turbine 24 and low pressure turbine 26 .
- Combusted primary air A P and secondary air A S are passed through augmentor duct 34 and into augmentation system 12 where a secondary combustion process can be carried out.
- Augmentation liner 36 prevents heat damage to augmentation system 12 and turbofan engine 10 .
- Exhausted air A Ex exits turbofan engine 10 through exhaust nozzle 28 .
- the adaptive cooling structure of the present invention can be used in combustor 22 or augmentation system 12 .
- adaptive cooling structure 40 such as augmentation liner 36 in augmentation system 12 or a heat shield in combustor 22 ( FIG. 1 ), is exposed directly to hot air A p .
- Adaptive cooling structure 40 includes liner 42 and mounting support 44 .
- Liner 42 is affixed to mounting support 44 by fastening means 46 such as threaded studs, bolts, rivets, welds, or other suitable fastening means.
- Liner 42 includes liner wall 48 with one or more film apertures 50 .
- Liner wall 48 has first surface 52 facing away from the mounting support 44 and second surface 54 facing towards mounting support 44 .
- Liner wall 48 may be made from a high temperature, cast, forged or sheet material such as nickel or cobalt for example.
- First surface 52 may also include one or more layers of thermal barrier coating (TBC) 56 , such as a metallic or ceramic material, for improved insulation from hot air A p .
- TBC thermal barrier coating
- Thermal gradient lines 58 depict the temperature differential across first surface 52 and indicate that hot spot location 60 is present in the area of liner 42 . Spallation of TBC layer 56 is also indicative of the presence of hot spot location 60 .
- Mounting support 44 includes one or more coolant apertures 62 .
- Coolant apertures 62 in mounting support 44 direct cooling air A C , such as pressurized air bled from compressor 18 or 20 ( FIG. 1 ), to second surface 54 of liner 42 .
- Coolant apertures 62 are perpendicular to the flow of hot air A p .
- coolant apertures 62 can be angled to the flow. Cooling air A C provides cooling to reduce the operating temperature of mounting support 44 as it flows through coolant apertures 62 . Cooling air A C exits coolant apertures 62 , flows between mounting support 44 and liner wall 48 , impinging on second surface 54 . Cooling air A C exits liner 42 through film apertures 50 in liner wall 48 , and provides film cooling of first surface 52 .
- liner 42 is porous instead of having film apertures 50 , and cooling air A C exits liner 42 through the pores.
- the present invention combines the benefits of both impingement cooling and film cooling and is particularly useful in parts such as combustor 22 and augmentation system 12 ( FIG. 1 ) where local hot spots develop.
- hot spot location 60 causes liner wall 48 to deflect away from mounting support 44 (as seen in FIG. 4 ).
- Impingement cooling has parameters which when engineered can provide an increased impingement rate upon deflection of liner wall 48 .
- the present invention configures these parameters to accommodate such deflections as ignoring these parameters results in a less efficient cooling structure.
- FIG. 3 is a cross-sectional view of adaptive cooling structure 40 taken at a non-hot spot location along line 3 - 3 of FIG. 2 .
- Liner 42 of adaptive cooling structure 40 includes mounting post 64 .
- Mounting post 64 with fastening means 46 is surrounded by spacer 66 and extends from second surface 54 of liner wall 48 through mounting support 44 .
- Nut 68 secures mounting post 64 to mounting support 44 via fastening means (threads) 46 .
- Spacer 66 such as a washer or other suitable spacer, creates chamber 70 between mounting support 44 and liner 42 for impingement cooling. Chamber 70 has distance L between mounting support 44 and liner 42 .
- Coolant apertures 62 have a circular cross section with diameter D. In other embodiments, coolant apertures 62 can have a non-circular cross section with effective diameter D.
- Adaptive cooling structure 40 is directly exposed to hot air Ap. Cooling air A C flows through coolant apertures 62 and enters chamber 70 , impinging on second surface 54 . Cooling air A C exits first surface 52 through film apertures 50 in liner wall 48 , forming a film. Film apertures 50 have a circular cross section, but can have a non-circular cross section or can be flared. Film apertures 50 are angled with the flow of hot air A P . In alternative embodiments, film apertures 50 can be at another angle or can be perpendicular to the flow. The location of coolant apertures 62 is staggered in relation to film apertures 50 . In alternative embodiments, the location of coolant apertures 62 can be aligned with film apertures 50 or completely independent of the location of film apertures 50 .
- a ratio L/D of distance L to diameter D of approximately three provides a preferred impingement heat transfer coefficient.
- distance L increases and ratio L/D increases as a result.
- the present invention is designed to accommodate the deformation by configuring adaptive cooling structure 40 with a ratio L/D lower than three.
- the preferred as-fabricated ratio L/D is in the range between approximately two and three, and more specifically 2.5. The configuration of the present invention thus results in increased impingement cooling effectiveness upon deformation in the hot spot, where it is most needed.
- FIG. 4 is a cross-sectional view of adaptive cooling structure 40 taken at a hot spot location along line 4 - 4 of FIG. 2 .
- Liner wall 48 is deflected away from mounting support 44 due to extreme heat caused by hot spot location 60 .
- Hot spot location 60 is exacerbated by an area of spalled TBC layer 56 .
- the deflection of liner wall 48 expanded chamber 70 , increasing distance L to L+ ⁇ L at hot spot location 60 and in turn increasing ratio L/D of distance L to diameter D of coolant apertures 62 .
- Cooling air A C flows through coolant apertures 62 and enters chamber 70 , impinging on second surface 54 . Cooling air A C exits first surface 52 through film apertures 50 in liner wall 48 , forming a film. Impingement effectiveness is increased at hot spot location 60 as a result of the deflection of liner 48 away from mounting support 44 .
- the fabrication of adaptive cooling structure 40 with a ratio L/D lower than the preferred ratio of three provides for increased impingement effectiveness when the deflection of liner wall 48 at hot spot location 60 increases distance L to L+ ⁇ L.
- the preferred increased ratio L/D resulting from the deflection of liner wall 48 is between three and 3.5, which results in a preferred impingement heat transfer coefficient.
- the increased ratio L/D ratio can be between approximately one and four or between two and four.
- FIG. 5 is a graph of ratio L/D versus impingement effectiveness H including preferred impingement effectiveness range 72 .
- the deflection of liner wall 48 in hot spot location 60 will increase the impingement effectiveness to range 72 .
- the deflection of liner wall 48 in hot spot location 60 would result in decreased impingement effectiveness range 76 .
- the present invention is specifically designed so the deflection of liner wall 48 results in ratio L/D in preferred impingement effectiveness range 72 .
- Impingement effectiveness range 72 can have L/D of between one and four, between two and four, or between 2.5 and 3.5.
- the preferred as-fabricated range 74 has ratio L/D of between approximately two and three, but can be anything less than three.
- decreased impingement effectiveness range 76 has ratio L/D of anything above four.
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Abstract
An adaptive cooling structure comprises a mounting support, a liner, and a spacer. The mounting support has a coolant aperture for directing cooling air through the support. The liner has a first surface facing away from the mounting support and a second surface facing towards the support. The liner is coupled to the mounting support, and the spacer is positioned between the support and the liner. The positioning of the spacer creates a chamber between the mounting support and the liner, thus allowing the cooling air to impinge on the second surface of the liner. The liner wall is configured to deflect away from the mounting support to expand the chamber, thus allowing the cooling air to further impinge on the second surface of the liner.
Description
- The present invention relates to cooling systems, and in particular, to a system and method for adaptive impingement cooling for use in hot environments such as those found in gas turbine engines.
- Gas turbine engines operate according to a continuous Brayton cycle where a pressurized air and fuel mixture is ignited in a combustor to produce a flowing stream of hot gas. The air is compressed, used for combustion, expands through a turbine, and finally exits the engine. Some gas turbine engines also include an augmentation system downstream of the turbine, where fuel is also introduced and ignited to increase thrust. Most often, the temperature of the primary air is higher than the melting temperatures of the materials that form the combustor, turbine, and augmentation system components. As a result, adequate cooling is integral to the function of gas turbine engines.
- It is common to combine the benefits of both impingement cooling and film cooling in gas turbine engines. This combination of impingement and film cooling is particularly useful in parts such as combustors and augmentation systems where local hot spots develop. Current practice is to design impingement cooling structures neglecting the deformation that occurs in local hot spots as the temperature in the hot spots increases. As a result, impingement cooling effectiveness decreases as the deformation develops, causing hot spots to become even hotter. Cooling effectiveness should be the highest at local hot spots.
- An adaptive cooling structure comprises a mounting support, a liner, and a spacer. The mounting support has a coolant aperture for directing cooling air through the support. The liner has a first surface facing away from the mounting support and a second surface facing towards the support. The liner is coupled to the mounting support, and the spacer is positioned between the support and the liner. The positioning of the spacer creates a chamber between the mounting support and the liner, thus allowing the cooling air to impinge on the second surface of the liner. The liner wall is configured to deflect away from the mounting support to expand the chamber, thus allowing the cooling air to further impinge on the second surface of the liner.
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FIG. 1 is a simplified cross-sectional view of an embodiment of a gas turbine engine which employs the adaptive impingement cooling system and method of the present invention. -
FIG. 2 is a partial isometric view of an embodiment of the adaptive cooling structure of the present invention. -
FIG. 3 is a cross-sectional view of the embodiment of the adaptive cooling structure inFIG. 2 at a non-hot spot location. -
FIG. 4 is a cross-sectional view of the embodiment of the adaptive cooling structure inFIG. 2 at a hot spot location. -
FIG. 5 is a graph showing preferred ranges of impingement effectiveness for designing the adaptive cooling structure of the present invention. -
FIG. 1 is a simplified cross-sectional view of mixedflow turbofan engine 10 which can employ the adaptive impingement cooling system and method of the present invention. Turbofanengine 10 includesaugmentation system 12,fan duct 14,drive fan 16,low pressure compressor 18,high pressure compressor 20,combustor 22,high pressure turbine 24,low pressure turbine 26, andexhaust nozzle 28.Drive fan 16 andlow pressure compressor 18 are driven bylow pressure turbine 26 withshaft 30.High pressure compressor 20 is driven byhigh pressure turbine 24 withshaft 32.High pressure compressor 20,combustor 22,high pressure turbine 24 andshaft 32 comprise the core ofturbofan engine 10.Augmentation system 12 includes augmenterduct 34 and augmenterliner 36. - Ambient air AAmbient enters
turbofan engine 10 atinlet 38 throughdrive fan 16. Drivefan 16 is rotated bylow pressure turbine 26 to accelerate AAmbient thereby producing a major portion of the thrust output ofturbofan engine 10. Accelerated AAmbient is divided into two streams of air: primary air AP and secondary air AS. Secondary air AS, also known as bypass air, passes intofan duct 14 where it passes on toaugmentation system 12. Primary air AP, also known as hot air, is a stream of air that is directed first intolow pressure compressor 18 and then intohigh pressure compressor 20. Pressurized primary air AP is then passed intocombustor 22 where it is mixed with a fuel supply and ignited to produce the high energy gases used to turnhigh pressure turbine 24 andlow pressure turbine 26. Combusted primary air AP and secondary air AS are passed throughaugmentor duct 34 and intoaugmentation system 12 where a secondary combustion process can be carried out.Augmentation liner 36 prevents heat damage toaugmentation system 12 andturbofan engine 10. Exhausted air AEx exitsturbofan engine 10 throughexhaust nozzle 28. The adaptive cooling structure of the present invention can be used incombustor 22 oraugmentation system 12. - Referring now to
FIG. 2 , adaptive cooling structure 40, such asaugmentation liner 36 inaugmentation system 12 or a heat shield in combustor 22 (FIG. 1 ), is exposed directly to hot air Ap. Adaptive cooling structure 40 includesliner 42 and mountingsupport 44.Liner 42 is affixed to mountingsupport 44 by fastening means 46 such as threaded studs, bolts, rivets, welds, or other suitable fastening means.Liner 42 includesliner wall 48 with one ormore film apertures 50.Liner wall 48 hasfirst surface 52 facing away from the mountingsupport 44 andsecond surface 54 facing towards mountingsupport 44.Liner wall 48 may be made from a high temperature, cast, forged or sheet material such as nickel or cobalt for example.First surface 52 may also include one or more layers of thermal barrier coating (TBC) 56, such as a metallic or ceramic material, for improved insulation from hot air Ap. Thermal gradient lines 58 depict the temperature differential acrossfirst surface 52 and indicate thathot spot location 60 is present in the area ofliner 42. Spallation ofTBC layer 56 is also indicative of the presence ofhot spot location 60. Mountingsupport 44 includes one ormore coolant apertures 62. -
Coolant apertures 62 in mountingsupport 44 direct cooling air AC, such as pressurized air bled fromcompressor 18 or 20 (FIG. 1 ), tosecond surface 54 ofliner 42.Coolant apertures 62 are perpendicular to the flow of hot air Ap. In an alternative embodiment,coolant apertures 62 can be angled to the flow. Cooling air AC provides cooling to reduce the operating temperature of mountingsupport 44 as it flows throughcoolant apertures 62. Cooling air AC exitscoolant apertures 62, flows between mountingsupport 44 andliner wall 48, impinging onsecond surface 54. Cooling air AC exitsliner 42 throughfilm apertures 50 inliner wall 48, and provides film cooling offirst surface 52. In an alternative embodiment,liner 42 is porous instead of havingfilm apertures 50, and cooling air AC exitsliner 42 through the pores. - The present invention combines the benefits of both impingement cooling and film cooling and is particularly useful in parts such as
combustor 22 and augmentation system 12 (FIG. 1 ) where local hot spots develop. Whenliner wall 48 of adaptive cooling structure 40 is exposed to hot air Ap,hot spot location 60 causesliner wall 48 to deflect away from mounting support 44 (as seen inFIG. 4 ). Impingement cooling has parameters which when engineered can provide an increased impingement rate upon deflection ofliner wall 48. Thus, the present invention configures these parameters to accommodate such deflections as ignoring these parameters results in a less efficient cooling structure. -
FIG. 3 is a cross-sectional view of adaptive cooling structure 40 taken at a non-hot spot location along line 3-3 ofFIG. 2 .Liner 42 of adaptive cooling structure 40 includes mountingpost 64. Mountingpost 64 with fastening means 46 is surrounded byspacer 66 and extends fromsecond surface 54 ofliner wall 48 through mountingsupport 44.Nut 68 secures mountingpost 64 to mountingsupport 44 via fastening means (threads) 46.Spacer 66, such as a washer or other suitable spacer, createschamber 70 between mountingsupport 44 andliner 42 for impingement cooling.Chamber 70 has distance L between mountingsupport 44 andliner 42.Coolant apertures 62 have a circular cross section with diameter D. In other embodiments,coolant apertures 62 can have a non-circular cross section with effective diameter D. - Adaptive cooling structure 40 is directly exposed to hot air Ap. Cooling air AC flows through
coolant apertures 62 and enterschamber 70, impinging onsecond surface 54. Cooling air AC exitsfirst surface 52 throughfilm apertures 50 inliner wall 48, forming a film.Film apertures 50 have a circular cross section, but can have a non-circular cross section or can be flared.Film apertures 50 are angled with the flow of hot air AP. In alternative embodiments,film apertures 50 can be at another angle or can be perpendicular to the flow. The location ofcoolant apertures 62 is staggered in relation tofilm apertures 50. In alternative embodiments, the location ofcoolant apertures 62 can be aligned withfilm apertures 50 or completely independent of the location offilm apertures 50. - In impingement cooling a ratio L/D of distance L to diameter D of approximately three provides a preferred impingement heat transfer coefficient. When
hot spot location 60 causesliner wall 48 to deflect away from mounting support 44 (as seen inFIG. 4 ), distance L increases and ratio L/D increases as a result. Thus, the present invention is designed to accommodate the deformation by configuring adaptive cooling structure 40 with a ratio L/D lower than three. For adaptive cooling structure 40, employing both impingement cooling and film cooling, the preferred as-fabricated ratio L/D is in the range between approximately two and three, and more specifically 2.5. The configuration of the present invention thus results in increased impingement cooling effectiveness upon deformation in the hot spot, where it is most needed. -
FIG. 4 is a cross-sectional view of adaptive cooling structure 40 taken at a hot spot location along line 4-4 ofFIG. 2 .Liner wall 48 is deflected away from mountingsupport 44 due to extreme heat caused byhot spot location 60.Hot spot location 60 is exacerbated by an area of spalledTBC layer 56. The deflection ofliner wall 48 expandedchamber 70, increasing distance L to L+ΔL athot spot location 60 and in turn increasing ratio L/D of distance L to diameter D ofcoolant apertures 62. - Cooling air AC flows through
coolant apertures 62 and enterschamber 70, impinging onsecond surface 54. Cooling air AC exitsfirst surface 52 throughfilm apertures 50 inliner wall 48, forming a film. Impingement effectiveness is increased athot spot location 60 as a result of the deflection ofliner 48 away from mountingsupport 44. As discussed in relation toFIG. 3 , the fabrication of adaptive cooling structure 40 with a ratio L/D lower than the preferred ratio of three provides for increased impingement effectiveness when the deflection ofliner wall 48 athot spot location 60 increases distance L to L+ΔL. Thus, the preferred increased ratio L/D resulting from the deflection ofliner wall 48 is between three and 3.5, which results in a preferred impingement heat transfer coefficient. In alternative embodiments, the increased ratio L/D ratio can be between approximately one and four or between two and four. -
FIG. 5 is a graph of ratio L/D versus impingement effectiveness H including preferredimpingement effectiveness range 72. If as-fabricated adaptive cooling structure 40 has ratio L/D inrange 74, less than approximately three, the deflection ofliner wall 48 inhot spot location 60 will increase the impingement effectiveness to range 72. If as-fabricated adaptive cooling structure 40 were to have ratio L/D equal to or greater than three, the deflection ofliner wall 48 inhot spot location 60 would result in decreasedimpingement effectiveness range 76. Thus, the present invention is specifically designed so the deflection ofliner wall 48 results in ratio L/D in preferredimpingement effectiveness range 72.Impingement effectiveness range 72 can have L/D of between one and four, between two and four, or between 2.5 and 3.5. As discussed in relation toFIG. 3 , the preferred as-fabricatedrange 74 has ratio L/D of between approximately two and three, but can be anything less than three. In one embodiment, decreasedimpingement effectiveness range 76 has ratio L/D of anything above four. - While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
Claims (21)
1. A structure for adaptive cooling comprising:
a mounting support having a coolant aperture for directing cooling air through the mounting support;
a liner coupled to the mounting support, including a wall having a first surface facing away from the mounting support and a second surface facing toward the mounting support;
a spacer positioned between the mounting support and the liner, the spacer creating a chamber between the mounting support and the liner, thus allowing the cooling air to impinge on the second surface of the liner; and
wherein the liner wall is configured to deflect away when exposed to hot air from the mounting support to expand the chamber, thus allowing the cooling air to further impinge on the second surface of the liner.
2. The structure of claim 1 , wherein the spacer positions the liner a distance away from the mounting support to provide impingement cooling at a first rate, and wherein the liner is configured to deflect an amount to increase the distance such that impingement cooling is provided at a second, greater rate.
3. The structure of claim 1 , wherein the liner permits the cooling air to pass through and exit the first surface, forming a film.
4. The structure of claim 1 , wherein the coolant aperture has a diameter D, the chamber has a distance L between the liner and the support that is less than three times the value of D, and the liner wall deflects away from the mounting support when exposed to hot air, increasing L to approximately three times the value of D.
5. The structure of claim 1 , wherein a mounting post with a threaded stud extends from the second surface of the liner wall and through the support, the mounting post is surrounded by a washer acting as the spacer between the support and the liner, and a nut secures the mounting post to the support.
6. The structure of claim 1 , wherein the first surface is a hot surface with a hot spot location, and the hot spot location causes the liner wall to deflect away from the mounting support.
7. The structure of claim 1 , wherein the liner is an impingement film cooled panel acting as a heat shield in a gas turbine combustor.
8. The structure of claim 1 , wherein the liner is an impingement film cooled liner in a gas turbine augmenter.
9. A method of adaptively cooling a liner coupled to a support with a spacer positioned between the liner and the support, the method comprising:
introducing cooling air into a coolant aperture in the support;
directing the cooling air into a chamber between the support and the liner and impinging the cooling air against the liner at a first rate;
deflecting the liner away from the mounting support, expanding the chamber; and
directing the cooling air into the chamber and further impinging the cooling air against the liner at a second rate.
10. The method of claim 9 , wherein the spacer positions the liner a distance away from the mounting support to provide impingement cooling at the first rate, and wherein the liner is configured to deflect an amount to increase the distance such that impingement cooling is provided at the second rate.
11. The method of claim 10 , wherein the second rate is greater than the first rate.
12. The method of claim 9 , wherein the coolant aperture has a diameter D, the chamber has a distance L between the liner and support that is less than three times the value of D, and the deflecting step causes the liner to deflect away from the mounting support, increasing L to between approximately one to four times the value of D.
13. The method of claim 10 , wherein the deflecting step causes the liner to deflect away from the mounting support, increasing L to between approximately two to four times the value of D.
14. The method of claim 10 , wherein the deflecting step causes the liner to deflect away from the mounting support, increasing L to approximately three times the value of D.
15. The method of claim 10 , wherein the chamber has a distance L between the liner and support that is between approximately two to three times the value of D, and the deflecting step causes the liner to deflect away from the mounting support, increasing L to between approximately two to four times the value of D.
16. The method of claim 13 , wherein the deflecting step causes the liner to deflect away from the mounting support, increasing L to between approximately 2.5 to 3.5 times the value of D.
17. The method of claim 13 , wherein the deflecting step causes the liner to deflect away from the mounting support, increasing L to approximately three times the value of D.
18. The method of claim 9 , and further comprising:
directing the cooling air to pass through the liner and exit the first surface, forming a film.
19. The method of claim 9 , wherein a hot spot location on the liner causes the deflecting step.
20. The method of claim 9 , wherein the liner is an impingement film cooled panel acting as a heat shield in a gas turbine combustor and the liner is exposed directly to hot air.
21. The method of claim 9 , wherein the liner is an impingement film cooled liner in a gas turbine augmenter and the liner is exposed directly to hot air.
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US13/174,166 US20130000309A1 (en) | 2011-06-30 | 2011-06-30 | System and method for adaptive impingement cooling |
EP12174426.2A EP2541147A3 (en) | 2011-06-30 | 2012-06-29 | System and method for adaptive impingement cooling |
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Cited By (13)
Publication number | Priority date | Publication date | Assignee | Title |
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US20110185738A1 (en) * | 2009-12-29 | 2011-08-04 | Bastnagel Philip M | Gas turbine engine component construction |
US20140204532A1 (en) * | 2013-01-21 | 2014-07-24 | Parker-Hannifin Corporation | Passively controlled smart microjet cooling array |
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US20180016918A1 (en) * | 2016-07-13 | 2018-01-18 | MTU Aero Engines AG | Shrouded blade of a gas turbine engine |
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Citations (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5131222A (en) * | 1990-11-28 | 1992-07-21 | The United States Of Americas As Represented By The Secretary Of The Air Force | Thermally valved cooling system for exhaust nozzle systems |
US5209059A (en) * | 1991-12-27 | 1993-05-11 | The United States Of America As Represented By The Secretary Of The Air Force | Active cooling apparatus for afterburners |
US7140185B2 (en) * | 2004-07-12 | 2006-11-28 | United Technologies Corporation | Heatshielded article |
US20100077764A1 (en) * | 2008-10-01 | 2010-04-01 | United Technologies Corporation | Structures with adaptive cooling |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6964170B2 (en) * | 2003-04-28 | 2005-11-15 | Pratt & Whitney Canada Corp. | Noise reducing combustor |
US8800298B2 (en) * | 2009-07-17 | 2014-08-12 | United Technologies Corporation | Washer with cooling passage for a turbine engine combustor |
-
2011
- 2011-06-30 US US13/174,166 patent/US20130000309A1/en not_active Abandoned
-
2012
- 2012-06-29 EP EP12174426.2A patent/EP2541147A3/en not_active Withdrawn
Patent Citations (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5131222A (en) * | 1990-11-28 | 1992-07-21 | The United States Of Americas As Represented By The Secretary Of The Air Force | Thermally valved cooling system for exhaust nozzle systems |
US5209059A (en) * | 1991-12-27 | 1993-05-11 | The United States Of America As Represented By The Secretary Of The Air Force | Active cooling apparatus for afterburners |
US7140185B2 (en) * | 2004-07-12 | 2006-11-28 | United Technologies Corporation | Heatshielded article |
US20100077764A1 (en) * | 2008-10-01 | 2010-04-01 | United Technologies Corporation | Structures with adaptive cooling |
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US20140204532A1 (en) * | 2013-01-21 | 2014-07-24 | Parker-Hannifin Corporation | Passively controlled smart microjet cooling array |
US9958160B2 (en) | 2013-02-06 | 2018-05-01 | United Technologies Corporation | Gas turbine engine component with upstream-directed cooling film holes |
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US20160370007A1 (en) * | 2014-05-09 | 2016-12-22 | United Technologies Corporation | Additively manufactured hotspot portion of a turbine engine component having heat resistant properties and method of manufacture |
US10935241B2 (en) * | 2014-05-09 | 2021-03-02 | Raytheon Technologies Corporation | Additively manufactured hotspot portion of a turbine engine component having heat resistant properties and method of manufacture |
JP2017525927A (en) * | 2014-08-26 | 2017-09-07 | シーメンス エナジー インコーポレイテッド | Film cooling hole array for an acoustic resonator in a gas turbine engine |
US10359194B2 (en) | 2014-08-26 | 2019-07-23 | Siemens Energy, Inc. | Film cooling hole arrangement for acoustic resonators in gas turbine engines |
US10132498B2 (en) * | 2015-01-20 | 2018-11-20 | United Technologies Corporation | Thermal barrier coating of a combustor dilution hole |
US20160209033A1 (en) * | 2015-01-20 | 2016-07-21 | United Technologies Corporation | Combustor dilution hole passive heat transfer control |
US20180016918A1 (en) * | 2016-07-13 | 2018-01-18 | MTU Aero Engines AG | Shrouded blade of a gas turbine engine |
US10544687B2 (en) * | 2016-07-13 | 2020-01-28 | MTU Aero Engines AG | Shrouded blade of a gas turbine engine |
DE102016222099A1 (en) * | 2016-11-10 | 2018-05-17 | Rolls-Royce Deutschland Ltd & Co Kg | Combustion chamber of a gas turbine |
Also Published As
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EP2541147A2 (en) | 2013-01-02 |
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