JP2010526274A5 - - Google Patents

Download PDF

Info

Publication number
JP2010526274A5
JP2010526274A5 JP2010506353A JP2010506353A JP2010526274A5 JP 2010526274 A5 JP2010526274 A5 JP 2010526274A5 JP 2010506353 A JP2010506353 A JP 2010506353A JP 2010506353 A JP2010506353 A JP 2010506353A JP 2010526274 A5 JP2010526274 A5 JP 2010526274A5
Authority
JP
Japan
Prior art keywords
diameter
liner
combustor
cooling
combustor liner
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.)
Pending
Application number
JP2010506353A
Other languages
Japanese (ja)
Other versions
JP2010526274A (en
Filing date
Publication date
Priority claimed from US11/743,126 external-priority patent/US20080271457A1/en
Application filed filed Critical
Publication of JP2010526274A publication Critical patent/JP2010526274A/en
Publication of JP2010526274A5 publication Critical patent/JP2010526274A5/ja
Pending legal-status Critical Current

Links

Claims (15)

ガスタービン燃焼器のライナであって、
（ａ）前記燃焼器の上流端部に隣接した第1の端部及び該燃焼器の下流端部に隣接した第２の端部を有しかつ高温側面、低温側面及びそれを貫通する中心軸線をさらに有するシェルと、
（ｂ）前記シェル内に形成されかつそれらを通して空気が流れて該シェルの高温側面に沿って冷却フィルムを形成するようになった複数の小径かつ密に間隔を置いて配置されたフィルム冷却孔と、を含み、
前記冷却孔が、前記シェルを貫通する一様でない直径を有する、
燃焼器ライナ。 A liner for a gas turbine combustor,
(A) a first end adjacent to the upstream end of the combustor and a second end adjacent to the downstream end of the combustor, and a hot side, a cold side and a central axis passing therethrough A shell further comprising:
(B) a plurality of small diameter and closely spaced film cooling holes formed in the shell and through which air flows to form a cooling film along the hot side of the shell; Including,
The cooling holes have non-uniform diameters through the shell;
Combustor liner. 前記冷却孔の各々が、
（ａ）前記シェルの低温側面に設置されかつ第１の直径を有する第１の開口と、
（ｂ）前記シェルの高温側面に設置されかつ第２の直径を有する第２の開口と、をさらに含み、
前記第２の開口の第２の直径が、前記第１の開口の第１の直径よりも大きい、
請求項１記載の燃焼器ライナ。 Each of the cooling holes is
(A) a first opening installed on a cold side of the shell and having a first diameter;
(B) a second opening disposed on the hot side of the shell and having a second diameter;
A second diameter of the second opening is greater than a first diameter of the first opening;
The combustor liner of claim 1. 前記第１の開口の第１の直径に対する前記第２の開口の第２の直径の比が、約３．０〜５．０である、請求項２記載の燃焼器ライナ。 The combustor liner of claim 2, wherein a ratio of a second diameter of the second opening to a first diameter of the first opening is about 3.0 to 5.0. 前記冷却孔の隣り合う第１の開口間の間隔が、該冷却孔の隣り合う第２の開口間の間隔よりも大きい、請求項２または３記載の燃焼器ライナ。 The combustor liner according to claim 2 or 3, wherein a distance between adjacent first openings of the cooling holes is larger than a distance between adjacent second openings of the cooling holes. 前記冷却孔の隣り合う第１の開口間の間隔が、前記第１の直径の約３．０〜６．０倍である、請求項２乃至４のいずれか１項記載の燃焼器ライナ。 The combustor liner according to any one of claims 2 to 4, wherein a spacing between adjacent first openings of the cooling holes is about 3.0 to 6.0 times the first diameter. 前記冷却孔の隣り合う第２の開口間の間隔が、前記第２の直径の約０．２〜０．７倍である、請求項２乃至５のいずれか１項記載の燃焼器ライナ。 The combustor liner of any one of claims 2 to 5, wherein a spacing between adjacent second openings of the cooling holes is about 0.2 to 0.7 times the second diameter. 各前記冷却孔の直径が、前記シェルの低温側面から該シェルの高温側面まで次第により大きくなる、請求項１乃至６のいずれか１項記載の燃焼器ライナ。 The combustor liner according to any one of claims 1 to 6, wherein the diameter of each cooling hole gradually increases from the cold side of the shell to the hot side of the shell. 各前記冷却孔が、ほぼ切頭円錐形状を有する、請求項７記載の燃焼器ライナ。 The combustor liner of claim 7, wherein each cooling hole has a generally frustoconical shape. 各前記冷却孔が、それを貫通する軸線に対して約１°〜約１５°の拡散角度を有する、請求項７または８記載の燃焼器ライナ。 The combustor liner of claim 7 or 8, wherein each cooling hole has a diffusion angle of about 1 ° to about 15 ° relative to an axis extending therethrough. 各前記冷却孔が、それを貫通する軸線に対して約３°〜約１０°の拡散角度を有する、請求項７乃至９のいずれか１項記載の燃焼器ライナ。 The combustor liner of any one of claims 7 to 9, wherein each cooling hole has a diffusion angle of about 3 ° to about 10 ° relative to an axis extending therethrough. 各前記冷却孔が、それを貫通する軸線に対して約５°〜約９°の拡散角度を有する、請求項７乃至１０のいずれか１項記載の燃焼器ライナ。 11. A combustor liner according to any one of claims 7 to 10, wherein each cooling hole has a diffusion angle of about 5 [deg.] To about 9 [deg.] Relative to an axis extending therethrough. 各前記冷却孔を貫通する軸線が、前記中心軸線に対して約１５°〜約３５°の範囲の軸方向角度で配向される、請求項１乃至１１のいずれか１項記載の燃焼器ライナ。 Axis extending through the respective said cooling holes, wherein are oriented in the axial direction an angle in the range of about 15 ° ~ about 35 ° with respect to the central axis, the combustor liner of any one of claims 1 to 11. 各前記冷却孔を貫通する軸線が、前記中心軸線に対して約３０°〜約６０°の範囲の円周方向角度で配向される、請求項１乃至１１のいずれか１項記載の燃焼器ライナ。 Axis extending through the respective said cooling holes, wherein is oriented in the circumferential direction an angle in the range of about 30 ° ~ about 60 ° with respect to the central axis, the combustor liner according to any one of claims 1 to 11 . それを貫通した一様でない直径を有する冷却孔をガスタービンエンジン燃焼器のライナ内に形成する方法であって、
（ａ）その形状がほぼ円錐形でありかつ前記ライナをほぼ貫通して延びる前記冷却孔の第１の部分を該ライナの高温側面から形成するステップと、
（ｂ）その直径がほぼ一様である前記冷却孔の第２の部分を該冷却孔の第１の部分から前記ライナの低温側面まで形成するステップと、を含む、
方法。 A method of forming a cooling hole having a non-uniform diameter therethrough in a liner of a gas turbine engine combustor, comprising:
(A) forming a first portion of the cooling hole having a generally conical shape and extending substantially through the liner from a hot side surface of the liner;
(B) forming a second portion of the cooling hole having a substantially uniform diameter from the first portion of the cooling hole to a cold side of the liner;
Method. 前記冷却孔の第１の部分が、前記ライナの高温側面から次第に減少する直径を有する、請求項１４記載の方法。 The method of claim 14 , wherein the first portion of the cooling hole has a diameter that gradually decreases from a hot side of the liner.
JP2010506353A 2007-05-01 2008-03-04 Cooling holes for gas turbine combustor liners with non-uniform diameters therethrough Pending JP2010526274A (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US11/743,126 US20080271457A1 (en) 2007-05-01 2007-05-01 Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough
PCT/US2008/055758 WO2008137201A1 (en) 2007-05-01 2008-03-04 Cooling holes for gas turbine combustor liner having a non-uniform diameter therethrough

Publications (2)

Publication Number Publication Date
JP2010526274A JP2010526274A (en) 2010-07-29
JP2010526274A5 true JP2010526274A5 (en) 2011-04-21

Family

ID=39645559

Family Applications (1)

Application Number Title Priority Date Filing Date
JP2010506353A Pending JP2010526274A (en) 2007-05-01 2008-03-04 Cooling holes for gas turbine combustor liners with non-uniform diameters therethrough

Country Status (5)

Country Link
US (1) US20080271457A1 (en)
EP (1) EP2153133A1 (en)
JP (1) JP2010526274A (en)
CA (1) CA2685342A1 (en)
WO (1) WO2008137201A1 (en)

Families Citing this family (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102009007164A1 (en) 2009-02-03 2010-08-12 Rolls-Royce Deutschland Ltd & Co Kg A method of forming a cooling air opening in a wall of a gas turbine combustor and combustor wall made by the method
US8307657B2 (en) * 2009-03-10 2012-11-13 General Electric Company Combustor liner cooling system
US20100263384A1 (en) * 2009-04-17 2010-10-21 Ronald James Chila Combustor cap with shaped effusion cooling holes
US8371814B2 (en) * 2009-06-24 2013-02-12 Honeywell International Inc. Turbine engine components
US8529193B2 (en) * 2009-11-25 2013-09-10 Honeywell International Inc. Gas turbine engine components with improved film cooling
US8628293B2 (en) 2010-06-17 2014-01-14 Honeywell International Inc. Gas turbine engine components with cooling hole trenches
US9052113B1 (en) 2011-06-06 2015-06-09 General Electric Company Combustor nozzle and method for modifying the combustor nozzle
US9650900B2 (en) 2012-05-07 2017-05-16 Honeywell International Inc. Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations
US9194585B2 (en) * 2012-10-04 2015-11-24 United Technologies Corporation Cooling for combustor liners with accelerating channels
US10113433B2 (en) 2012-10-04 2018-10-30 Honeywell International Inc. Gas turbine engine components with lateral and forward sweep film cooling holes
US9309809B2 (en) * 2013-01-23 2016-04-12 General Electric Company Effusion plate using additive manufacturing methods
US9939154B2 (en) * 2013-02-14 2018-04-10 United Technologies Corporation Combustor liners with U-shaped cooling channels
JP6438000B2 (en) * 2013-03-15 2018-12-19 プレジデント アンド フェローズ オブ ハーバード カレッジ Low porosity auxetic sheet
US11112115B2 (en) * 2013-08-30 2021-09-07 Raytheon Technologies Corporation Contoured dilution passages for gas turbine engine combustor
EP3066388B1 (en) * 2013-11-04 2024-04-10 RTX Corporation Turbine engine combustor heat shield with multi-angled cooling apertures
EP3066391B1 (en) * 2013-11-05 2019-01-16 United Technologies Corporation Cooled combustor floatwall panel
WO2016032434A1 (en) * 2014-08-26 2016-03-03 Siemens Energy, Inc. Film cooling hole arrangement for acoustic resonators in gas turbine engines
GB201419327D0 (en) 2014-10-30 2014-12-17 Rolls Royce Plc A cooled component
US20160258623A1 (en) * 2015-03-05 2016-09-08 United Technologies Corporation Combustor and heat shield configurations for a gas turbine engine
US11021965B2 (en) 2016-05-19 2021-06-01 Honeywell International Inc. Engine components with cooling holes having tailored metering and diffuser portions
US20180266687A1 (en) * 2017-03-16 2018-09-20 General Electric Company Reducing film scrubbing in a combustor
US11149948B2 (en) * 2017-08-21 2021-10-19 General Electric Company Fuel nozzle with angled main injection ports and radial main injection ports
US11248791B2 (en) 2018-02-06 2022-02-15 Raytheon Technologies Corporation Pull-plane effusion combustor panel
US10830435B2 (en) 2018-02-06 2020-11-10 Raytheon Technologies Corporation Diffusing hole for rail effusion
US11009230B2 (en) 2018-02-06 2021-05-18 Raytheon Technologies Corporation Undercut combustor panel rail
US11022307B2 (en) * 2018-02-22 2021-06-01 Raytheon Technology Corporation Gas turbine combustor heat shield panel having multi-direction hole for rail effusion cooling
US11306659B2 (en) 2019-05-28 2022-04-19 Honeywell International Inc. Plug resistant effusion holes for gas turbine engine
US11371701B1 (en) 2021-02-03 2022-06-28 General Electric Company Combustor for a gas turbine engine
US11774098B2 (en) 2021-06-07 2023-10-03 General Electric Company Combustor for a gas turbine engine
US11885495B2 (en) 2021-06-07 2024-01-30 General Electric Company Combustor for a gas turbine engine including a liner having a looped feature
US11959643B2 (en) 2021-06-07 2024-04-16 General Electric Company Combustor for a gas turbine engine
CN115493163B (en) * 2022-09-06 2024-02-20 清华大学 Combustion chamber flame tube and high-efficiency cooling method thereof

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2221979B (en) * 1988-08-17 1992-03-25 Rolls Royce Plc A combustion chamber for a gas turbine engine
US5660525A (en) * 1992-10-29 1997-08-26 General Electric Company Film cooled slotted wall
US5528904A (en) * 1994-02-28 1996-06-25 Jones; Charles R. Coated hot gas duct liner
JPH09119322A (en) * 1995-10-27 1997-05-06 Ishikawajima Harima Heavy Ind Co Ltd Cooling liner of aircraft engine
US6383602B1 (en) * 1996-12-23 2002-05-07 General Electric Company Method for improving the cooling effectiveness of a gaseous coolant stream which flows through a substrate, and related articles of manufacture
DE59808269D1 (en) * 1998-03-23 2003-06-12 Alstom Switzerland Ltd Film cooling hole
CA2288557C (en) * 1998-11-12 2007-02-06 Mitsubishi Heavy Industries, Ltd. Gas turbine combustor cooling structure
US6408629B1 (en) * 2000-10-03 2002-06-25 General Electric Company Combustor liner having preferentially angled cooling holes
US7093439B2 (en) * 2002-05-16 2006-08-22 United Technologies Corporation Heat shield panels for use in a combustor for a gas turbine engine
US7328580B2 (en) * 2004-06-23 2008-02-12 General Electric Company Chevron film cooled wall
JP4898253B2 (en) * 2005-03-30 2012-03-14 三菱重工業株式会社 High temperature components for gas turbines
EP1712739A1 (en) * 2005-04-12 2006-10-18 Siemens Aktiengesellschaft Component with film cooling hole

Similar Documents

Publication Publication Date Title
JP2010526274A5 (en)
JP2011220669A5 (en)
US20120024985A1 (en) Integrated fuel nozzle and inlet flow conditioner and related method
US9810081B2 (en) Cooled conduit for conveying combustion gases
EP2101041B1 (en) Cooling air manifold splash plate for a gas turbine engine
WO2009028475A1 (en) Gas turbine combustor
JP2012149869A5 (en)
US9046269B2 (en) Impingement cooling device
JP2006242561A (en) Combustor liner assembly and combustor assembly
JP2011058792A5 (en)
JP2011157963A5 (en)
JP2010526274A (en) Cooling holes for gas turbine combustor liners with non-uniform diameters therethrough
RU2718371C2 (en) Annular wall of combustion chamber with optimized cooling
EP1811231A3 (en) Segmented effusion cooled gas turbine engine combustor
JP2010085083A5 (en)
EP2163819A3 (en) Combustor, method of supplying fuel to same, and method of modifying same
CA2587060A1 (en) Combustor with improved swirl
JP2009168439A (en) Swozzle (fuel nozzle) design for gas turbine combustor
JP2012149881A5 (en)
JP2011137629A5 (en)
JP2011094951A (en) Apparatus for conditioning airflow passing through nozzle
JP2011094951A5 (en)
EP2196634A3 (en) Cavity ventilation
JP2006242184A5 (en)
JP2013513777A (en) Turbine engine combustion chamber