EP2541147A2 - Procédé et système de refroidissement adaptatif par impact - Google Patents

Procédé et système de refroidissement adaptatif par impact Download PDF

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Publication number
EP2541147A2
EP2541147A2 EP12174426A EP12174426A EP2541147A2 EP 2541147 A2 EP2541147 A2 EP 2541147A2 EP 12174426 A EP12174426 A EP 12174426A EP 12174426 A EP12174426 A EP 12174426A EP 2541147 A2 EP2541147 A2 EP 2541147A2
Authority
EP
European Patent Office
Prior art keywords
liner
support
mounting support
cooling
value
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.)
Withdrawn
Application number
EP12174426A
Other languages
German (de)
English (en)
Other versions
EP2541147A3 (fr
Inventor
James A. Dierberger
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2541147A2 publication Critical patent/EP2541147A2/fr
Publication of EP2541147A3 publication Critical patent/EP2541147A3/fr
Withdrawn legal-status Critical Current

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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
    • 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
    • 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
    • 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
    • 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 Ambieni is divided into two streams of air: primary air A P and secondary air As.
  • Secondary air As 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 Ap 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 . In an alternative embodiment, 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. In an alternative embodiment, 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 A P .
  • 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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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP12174426.2A 2011-06-30 2012-06-29 Procédé et système de refroidissement adaptatif par impact Withdrawn EP2541147A3 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/174,166 US20130000309A1 (en) 2011-06-30 2011-06-30 System and method for adaptive impingement cooling

Publications (2)

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EP2541147A2 true EP2541147A2 (fr) 2013-01-02
EP2541147A3 EP2541147A3 (fr) 2017-11-01

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Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014126641A1 (fr) * 2013-02-14 2014-08-21 United Technologies Corporation Ensemble support de chemise de protection thermique adaptatif destiné à des moteurs de turbine à gaz
WO2015054244A1 (fr) 2013-10-07 2015-04-16 United Technologies Corporation Paroi de dispositif de combustion assemblée par soudage pour un moteur à turbine
WO2015057272A1 (fr) * 2013-10-18 2015-04-23 United Technologies Corporation Paroi de chambre de combustion ayant un ou plusieurs éléments de refroidissement dans une cavité de refroidissement
WO2015084444A1 (fr) 2013-12-06 2015-06-11 United Technologies Corporation Interfaces d'ensemble paroi de turbine à gaz

Families Citing this family (11)

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Publication number Priority date Publication date Assignee Title
US9341118B2 (en) * 2009-12-29 2016-05-17 Rolls-Royce Corporation Various layered gas turbine engine component constructions
US9247672B2 (en) * 2013-01-21 2016-01-26 Parker-Hannifin Corporation Passively controlled smart microjet cooling array
WO2014123850A1 (fr) * 2013-02-06 2014-08-14 United Technologies Corporation Composant de turbine à gaz avec trous de film de refroidissement orientés vers l'amont
EP2954261B1 (fr) 2013-02-08 2020-03-04 United Technologies Corporation Chambre de combustion de turbine à gaz
US10386066B2 (en) * 2013-11-22 2019-08-20 United Technologies Corpoation Turbine engine multi-walled structure with cooling element(s)
US10344979B2 (en) * 2014-01-30 2019-07-09 United Technologies Corporation Cooling flow for leading panel in a gas turbine engine combustor
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
CN107076416B (zh) 2014-08-26 2020-05-19 西门子能源公司 用于燃气涡轮发动机中的声共振器的薄膜冷却孔装置
US10132498B2 (en) * 2015-01-20 2018-11-20 United Technologies Corporation Thermal barrier coating of a combustor dilution hole
EP3269932A1 (fr) * 2016-07-13 2018-01-17 MTU Aero Engines GmbH Aube carénée pour turbine à gaz
DE102016222099A1 (de) * 2016-11-10 2018-05-17 Rolls-Royce Deutschland Ltd & Co Kg Brennkammer einer Gasturbine

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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
US6964170B2 (en) * 2003-04-28 2005-11-15 Pratt & Whitney Canada Corp. Noise reducing combustor
US7140185B2 (en) * 2004-07-12 2006-11-28 United Technologies Corporation Heatshielded article
US9587832B2 (en) * 2008-10-01 2017-03-07 United Technologies Corporation Structures with adaptive cooling
US8800298B2 (en) * 2009-07-17 2014-08-12 United Technologies Corporation Washer with cooling passage for a turbine engine combustor

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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014126641A1 (fr) * 2013-02-14 2014-08-21 United Technologies Corporation Ensemble support de chemise de protection thermique adaptatif destiné à des moteurs de turbine à gaz
US10077681B2 (en) 2013-02-14 2018-09-18 United Technologies Corporation Compliant heat shield liner hanger assembly for gas turbine engines
WO2015054244A1 (fr) 2013-10-07 2015-04-16 United Technologies Corporation Paroi de dispositif de combustion assemblée par soudage pour un moteur à turbine
EP3055530A4 (fr) * 2013-10-07 2016-11-09 United Technologies Corp Paroi de dispositif de combustion assemblée par soudage pour un moteur à turbine
WO2015057272A1 (fr) * 2013-10-18 2015-04-23 United Technologies Corporation Paroi de chambre de combustion ayant un ou plusieurs éléments de refroidissement dans une cavité de refroidissement
WO2015084444A1 (fr) 2013-12-06 2015-06-11 United Technologies Corporation Interfaces d'ensemble paroi de turbine à gaz
EP3077729A4 (fr) * 2013-12-06 2017-01-11 United Technologies Corporation Interfaces d'ensemble paroi de turbine à gaz

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Publication number Publication date
EP2541147A3 (fr) 2017-11-01
US20130000309A1 (en) 2013-01-03

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