EP3385622A1 - Gasturbinenbrennraumtülle - Google Patents

Gasturbinenbrennraumtülle Download PDF

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Publication number
EP3385622A1
EP3385622A1 EP18165211.6A EP18165211A EP3385622A1 EP 3385622 A1 EP3385622 A1 EP 3385622A1 EP 18165211 A EP18165211 A EP 18165211A EP 3385622 A1 EP3385622 A1 EP 3385622A1
Authority
EP
European Patent Office
Prior art keywords
combustor
grommet
panel
liner
radius
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Withdrawn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
EP18165211.6A
Other languages
English (en)
French (fr)
Inventor
San Quach
Robert M. Sonntag
Donald W. Peters
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 EP3385622A1 publication Critical patent/EP3385622A1/de
Withdrawn legal-status Critical Current

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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
    • F23R3/002—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
    • F23R3/06—Arrangement of apertures along the flame tube

Definitions

  • the present disclosure relates to grommets for gas turbine engines, and more particularly to grommets for use with combustors for gas turbine engines.
  • Dilution holes and grommets can be utilized within combustors of gas turbine engines to condition combustion gases prior to entering the turbine for reliable operation and performance.
  • the dilution hole grommets may experience high temperatures, oxidize, and contain features that reduce flow.
  • a combustor panel assembly for use with a combustor includes a combustor panel with a panel dilution hole formed therethrough, a combustor liner spaced apart from the combustor panel, wherein the combustor liner includes a liner dilution hole formed therethrough, wherein at least one of the combustor panel and the combustor liner is formed from a first material, and a grommet with a first end with a first radius, a second end with a second radius, and a continuous smooth surface between the first end and the second end, wherein the first radius is larger than the second radius, the grommet defining a flow path between the panel dilution hole and the liner dilution hole, and the grommet is formed from a second material.
  • grommet has a frustroconical shape.
  • further embodiments could include that the second material is a ceramic matric composite.
  • further embodiments could include that the second material is a molybdenum alloy.
  • further embodiments could include that the second material is a nickel alloy.
  • further embodiments could include that the second material is a monolithic ceramic.
  • further embodiments could include that the grommet is affixed to the combustor panel.
  • further embodiments could include that the grommet is affixed to the combustor panel via a fastener.
  • further embodiments could include that the grommet is affixed to the combustor panel via a brazed joint.
  • further embodiments could include that the grommet is affixed to the combustor panel via a welded joint.
  • further embodiments could include that the grommet is affixed to the combustor panel via a threaded connection.
  • a combustor includes a combustor panel with a panel dilution hole formed therethrough, a combustor liner spaced apart from the combustor panel, wherein the combustor liner includes a liner dilution hole formed therethrough, wherein at least one of the combustor panel and the combustor liner is formed from a first material, and a grommet with a first end with a first radius, a second end with a second radius, and a continuous smooth surface between the first end and the second end, wherein the first radius is larger than the second radius, the grommet defining a flow path between the panel dilution hole and the liner dilution hole, and the grommet is formed from a second material.
  • grommet has a frustroconical shape.
  • further embodiments could include that the second material is a ceramic matric composite.
  • further embodiments could include that the second material is a molybdenum alloy.
  • further embodiments could include that the second material is a nickel alloy.
  • a gas turbine engine includes a combustor, including a combustor panel with a panel dilution hole formed therethrough, a combustor liner spaced apart from the combustor panel, wherein the combustor liner includes a liner dilution hole formed therethrough, wherein at least one of the combustor panel and the combustor liner is formed from a first material, and a grommet with a first end with a first radius, a second end with a second radius, and a continuous smooth surface between the first end and the second end, wherein the first radius is larger than the second radius, the grommet defining a flow path between the panel dilution hole and the liner dilution hole, and the grommet is formed from a second material.
  • grommet has a frustroconical shape.
  • further embodiments could include that the second material is a ceramic matric composite.
  • further embodiments could include that the second material is a molybdenum alloy.
  • Embodiments provide a grommet for use with a combustor panel assembly.
  • the grommet provides a low restriction flow path for conditioning airflow while adding high temperature resistant material to the combustor panel assembly to allow for reliable operation and desired performance.
  • the gas turbine engine 20 includes a fan section 22, a compressor section 24, a combustor section 26, and a turbine section 28 disposed about a longitudinal axis A.
  • the fan section 22 drives air along a bypass flow path B that may bypass the compressor section 24, the combustor section 26, and the turbine section 28.
  • the compressor section 24 draws air in along a core flow path C where air is compressed by the compressor section 24 and is provided to or communicated to the combustor section 26.
  • the compressed air is heated by the combustor section 26 to generate a high pressure exhaust gas stream that expands through the turbine section 28.
  • the turbine section 28 extracts energy from the high pressure exhaust gas stream to drive the fan section 22 and the compressor section 24.
  • the gas turbine engine 20 further includes a low-speed spool 30 and a high-speed spool 32 that are configured to rotate the fan section 22, the compressor section 24, and the turbine section 28 about the longitudinal axis A.
  • the low-speed spool 30 may connect a fan 48 of the fan section 22 and a low-pressure compressor portion 44 of the compressor section 24 to a low-pressure turbine portion 46 of the turbine section 28.
  • the turbine section 28 can include a rotating disc assembly.
  • the high-speed spool 32 may connect a high pressure compressor portion 52 of the compressor section 24 and a high pressure turbine portion 54 of the turbine section 28.
  • the fan 48 includes a fan rotor or fan hub that carries a fan blade 42. The fan blade 42 radially extends from the fan hub.
  • the combustor section 26 can have operating temperatures that exceed the melting point of the materials forming the combustor section 26 components.
  • the combustor section 16 can include dilution holes to condition combustion air.
  • the combustor section 16 includes a combustor panel assembly 17 with a combustor liner 60 and a combustor panel 62.
  • the combustor liner 60 and the combustor panel 62 can be formed from any suitable material.
  • the combustor panel 62 defines an annular shaped combustion chamber.
  • the combustor liner 60 is spaced apart from the combustor panel 62 to form an air passage between the combustor panel 62 and the combustor liner 60 to provide conditioning air.
  • the combustor panel 62 and the combustor liner 60 can include dilution holes to condition combustion air flow.
  • the combustor liner 60 includes liner dilution holes 64 and the combustor panel 62 includes panel dilution holes 66.
  • the air exiting the compressor section 14 of the gas turbine engine 10 is typically split or bifurcated, with a portion of the compressed air being used for combustion, and a portion of the compressed air being directed to the liner dilution holes 64 and the panel dilution holes 66 for conditioning.
  • the liner dilution holes 64 and the panel dilution holes 66 can receive conditioning airflow and provide conditioning airflow to combustor components.
  • the panel dilution holes 66 receive airflow from the liner dilution holes 64.
  • the combustor panel 62 includes an alignment stud 68.
  • the combustor liner 60 includes an alignment hole 69.
  • the alignment hole 69 can receive the alignment stud 68 to facilitate alignment and installation. Further, the use of the alignment hole 69 with the alignment stud 68 can allow for the liner dilution holes 64 to be aligned with the panel dilution holes 66.
  • a grommet 70 for use with the combustor panel assembly 17 is shown.
  • the grommet 70 includes an upper radius 72, a lower radius 74, an inner surface 76 and an outer surface 78.
  • the grommet 70 can be disposed within the liner dilution hole 64 and the panel dilution hole 66 to provide a smooth continuous flow path for the conditioning air flow.
  • the use of the grommet 70 can minimize flow restrictions for conditioning airflow while providing high-temperature resistance within the combustor section 16.
  • the grommet 70 is formed from a different material than the combustor panel 62 and the combustor liner 60.
  • the grommet 70 is formed from a ceramic matrix composite.
  • the grommet 70 is formed from monolithic ceramics.
  • the grommet 70 is formed from molybdenum alloys.
  • the grommet 70 is formed from nickel alloys.
  • the material of grommet 70 is selected to prevent melting or cracking of the grommet 70. Further, in certain embodiments, the materials of the grommet 70 are selected to have a low thermal mass to minimize stored thermal energy. Further, the material of grommet 70 can prevent the formation of oxidation.
  • the grommet 70 is generally shaped to smoothly direct air from the outside the combustor liner 60, through the liner dilution hole 64 and through the panel dilution hole 66.
  • the grommet 70 is fastened to the combustor panel 62 via the panel dilution hole 66. In certain embodiments, a pressure differential keeps the grommet 70 attached to the combustor panel 62. In other embodiments, the grommet 70 is fastened to the combustor liner 60 or a combination of the combustor panel 62 and the combustor liner 60. In certain embodiments, the grommet 70 is fastened via fasteners, such as nuts, bolts, etc., brazing, welding, or a threaded connection.
  • the geometry of the grommet 70 reduces flow restriction for conditioning flows entering the liner dilution hole 64 and the panel dilution hole 66, allowing for less work to be performed by the compressor of the engine.
  • the grommet 70 generally has a frustroconical shape, with an upper radius 72 and a lower radius 74. In the illustrated embodiment, the grommet 70 converges from the larger upper radius 72 to the smaller lower radius 74.
  • the inner surface 76 provides a continuous, uninterrupted flow path that transitions from the larger upper radius 72 to the smaller lower radius 74.
  • the outer surface 78 can interface with the combustor liner 60, the liner dilution hole 64, the combustor panel 62, and the panel dilution hole 66.
  • the materials and geometry of the grommet 70 allow for more efficient engine operation, increased durability, higher temperature operation, prevents thermal and mechanical failure, and allows for increased "time on wing" metrics.

Landscapes

  • 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)
EP18165211.6A 2017-04-03 2018-03-29 Gasturbinenbrennraumtülle Withdrawn EP3385622A1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US15/477,926 US20180283695A1 (en) 2017-04-03 2017-04-03 Combustion panel grommet

Publications (1)

Publication Number Publication Date
EP3385622A1 true EP3385622A1 (de) 2018-10-10

Family

ID=61868345

Family Applications (1)

Application Number Title Priority Date Filing Date
EP18165211.6A Withdrawn EP3385622A1 (de) 2017-04-03 2018-03-29 Gasturbinenbrennraumtülle

Country Status (2)

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US (1) US20180283695A1 (de)
EP (1) EP3385622A1 (de)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3039340B1 (de) * 2013-08-30 2018-11-28 United Technologies Corporation Vena-contracta-verwirbelungs-verdünnungspassagen für eine gasturbinenmotor-brennkammer
US20180299126A1 (en) * 2017-04-18 2018-10-18 United Technologies Corporation Combustor liner panel end rail
US20180306113A1 (en) * 2017-04-19 2018-10-25 United Technologies Corporation Combustor liner panel end rail matching heat transfer features
GB201902693D0 (en) * 2019-02-28 2019-04-17 Rolls Royce Plc Combustion liner and gas turbine engine comprising a combustion liner

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3594109A (en) * 1968-07-27 1971-07-20 Leyland Gass Turbines Ltd Flame tube
EP2738470A1 (de) * 2012-11-28 2014-06-04 Rolls-Royce Deutschland Ltd & Co KG Schindelbefestigungsanordnung einer Gasturbinenbrennkammer
WO2015030927A1 (en) * 2013-08-30 2015-03-05 United Technologies Corporation Contoured dilution passages for a gas turbine engine combustor

Family Cites Families (30)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1278590A (en) * 1968-09-20 1972-06-21 Lucas Industries Ltd Combustion chambers for gas turbine engines
US3545202A (en) * 1969-04-02 1970-12-08 United Aircraft Corp Wall structure and combustion holes for a gas turbine engine
US3981142A (en) * 1974-04-01 1976-09-21 General Motors Corporation Ceramic combustion liner
US4155680A (en) * 1977-02-14 1979-05-22 General Electric Company Compressor protection means
US4132066A (en) * 1977-09-23 1979-01-02 United Technologies Corporation Combustor liner for gas turbine engine
US4887432A (en) * 1988-10-07 1989-12-19 Westinghouse Electric Corp. Gas turbine combustion chamber with air scoops
US5291733A (en) * 1993-02-08 1994-03-08 General Electric Company Liner mounting assembly
US6351949B1 (en) * 1999-09-03 2002-03-05 Allison Advanced Development Company Interchangeable combustor chute
FR2825784B1 (fr) * 2001-06-06 2003-08-29 Snecma Moteurs Accrochage de chambre de combustion cmc de turbomachine utilisant les trous de dilution
US6904757B2 (en) * 2002-12-20 2005-06-14 General Electric Company Mounting assembly for the forward end of a ceramic matrix composite liner in a gas turbine engine combustor
US7153054B2 (en) * 2004-05-20 2006-12-26 United Technologies Corporation Fastener assembly for attaching a non-metal component to a metal component
GB2444736B (en) * 2006-12-12 2009-06-03 Rolls Royce Plc Combustion Chamber Air Inlet
US8387397B2 (en) * 2009-01-27 2013-03-05 General Electric Company Flow conditioner for use in gas turbine component in which combustion occurs
US20100212324A1 (en) * 2009-02-26 2010-08-26 Honeywell International Inc. Dual walled combustors with impingement cooled igniters
US20100242483A1 (en) * 2009-03-30 2010-09-30 United Technologies Corporation Combustor for gas turbine engine
US8397511B2 (en) * 2009-05-19 2013-03-19 General Electric Company System and method for cooling a wall of a gas turbine combustor
US8407892B2 (en) * 2011-08-05 2013-04-02 General Electric Company Methods relating to integrating late lean injection into combustion turbine engines
US9097424B2 (en) * 2012-03-12 2015-08-04 General Electric Company System for supplying a fuel and working fluid mixture to a combustor
US9360215B2 (en) * 2012-04-02 2016-06-07 United Technologies Corporation Combustor having a beveled grommet
DE102012015449A1 (de) * 2012-08-03 2014-02-20 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit Mischluftöffnungen und Luftleitelementen in modularer Bauweise
EP3039340B1 (de) * 2013-08-30 2018-11-28 United Technologies Corporation Vena-contracta-verwirbelungs-verdünnungspassagen für eine gasturbinenmotor-brennkammer
EP4353856A2 (de) * 2013-12-27 2024-04-17 Raytheon Technologies Corporation Hochfeste nickelknetlegierung mit hoher wärmeleitfähigkeit
US10151486B2 (en) * 2014-01-03 2018-12-11 United Technologies Corporation Cooled grommet for a combustor wall assembly
EP2957833B1 (de) * 2014-06-17 2018-10-24 Rolls-Royce Corporation Brennkammeranordnung mit rinnen
US20160115874A1 (en) * 2014-10-28 2016-04-28 Solar Turbines Incorporated Liner grommet assembly
US10612781B2 (en) * 2014-11-07 2020-04-07 United Technologies Corporation Combustor wall aperture body with cooling circuit
US20160209035A1 (en) * 2015-01-16 2016-07-21 Solar Turbines Incorporated Combustion hole insert with integrated film restarter
US20170059159A1 (en) * 2015-08-25 2017-03-02 Rolls-Royce Corporation Cmc combustor shell with integral chutes
US10520197B2 (en) * 2017-06-01 2019-12-31 General Electric Company Single cavity trapped vortex combustor with CMC inner and outer liners
US20190024895A1 (en) * 2017-07-18 2019-01-24 General Electric Company Combustor dilution structure for gas turbine engine

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3594109A (en) * 1968-07-27 1971-07-20 Leyland Gass Turbines Ltd Flame tube
EP2738470A1 (de) * 2012-11-28 2014-06-04 Rolls-Royce Deutschland Ltd & Co KG Schindelbefestigungsanordnung einer Gasturbinenbrennkammer
WO2015030927A1 (en) * 2013-08-30 2015-03-05 United Technologies Corporation Contoured dilution passages for a gas turbine engine combustor

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