US3593518A - Combustion chambers for gas turbine engines - Google Patents

Combustion chambers for gas turbine engines Download PDF

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Publication number
US3593518A
US3593518A US858770A US3593518DA US3593518A US 3593518 A US3593518 A US 3593518A US 858770 A US858770 A US 858770A US 3593518D A US3593518D A US 3593518DA US 3593518 A US3593518 A US 3593518A
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nozzles
air
combustion chamber
air inlet
inlet
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Expired - Lifetime
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US858770A
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Alan Joseph Gerrard
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ZF International UK Ltd
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Lucas Industries Ltd
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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
    • F23R3/06—Arrangement of apertures along the flame tube
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—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
    • 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
    • F23R3/26—Controlling the air flow

Definitions

  • a combustion chamber for a gas turbine engine has a primary air inlet and a number of additional air'nozzles at intervals along the chamber. Each of the air inlets has one or more holes in the inlet wall, so that air may be injected into the inlets or extracted therefrom. Air injection effectively reduces the inlet areas and extraction increases the inlet areas.
  • each inlet or group of inlets, may thus be controlled, the arrangement being such that, irrespective of the proportion of the total airflow which enters each inlet, the resistance to airflow through the combustion chamber does not vary.
  • the holes in the inlet walls may be tangential to create a vortex within the inlet.
  • This invention relates to combustion chambers for gas turbine engines and hasas an object to provide a combustion chamber in a convenient form.
  • a combustion chamber in accordance with the invention has a primary air inlet duct, secondary air inlet nozzles and dilution air inlet nozzles, each air entry point incorporating a fluidic airflow control device arranged so that the proportions of the total airflow through the respective air entry points can be varied without variation of the total resistance of the combustion chamber to airflow therethrough.
  • FIG. 1 is a fragmentary section through an annular combustion chamber incorporating an example of the invention
  • FIG. 2 is a fragmentary enlargement of the part ringed in FIG. 1 and FIGS. 3 and 4 show, somewhat diagrammatically, two views of a part of an alternative embodiment.
  • the combustion chamber has an inner wall and an outer wall 11 shaped tojoin at a circular leading end in which primary air inlet ducts 12 are formed.
  • Each primary air inlet duct has a swirler 13 which incorporates control jet holes 14 through which high-pressure air can be injected.
  • These holes 14 are arranged around the outer periphery of the swirler 13 so that when compressed air is injected the effective area of the swirler is reduced fluidically to reduce the airflow through the primary air inlet ducts. Conversely, when air is bled off through the holes 14 the effective area of the swirler is increased fluidically to increase the airflow through the primary inlet duct. 7
  • the walls 10, ll of the combustion chamber adjacent the inlet end thereof are formed with secondary air inlet nozzles 15 constituted by inwardly directed flanges on the walls.
  • an annular dished member 16 In association with the nozzles 15 is an annular dished member 16 defining an annular air chamber 17. Drillings are formed in the flanges which open into the nozzles 15 on the upstream sides thereof.
  • jets will issue into the nozzles 15 and fluidically restrict flow of air into the combustion chamber.
  • air is drawn from the chambers 17 the effective area of the nozzles 15 will be increased so that there will be increased airflow through the nozzles 15.
  • the walls 10, 11 Downstream of the nozzles 15 the walls 10, 11 have dilution air inlet nozzles 18 similar to the nozzles 15. These nozzles 18 similarly have associated therewith fluidic flow control devices constituted by drillings in the nozzle flanges opening into chambers 19 to which air is supplied to decrease the effective areas of nozzles 18 and from which air is drawn to increase the effective area.
  • the fluidic devices mentioned above are actuated to decrease flow through the primary and secondary air nozzles to a minimum value, whilst air is drawn from the chambers 19 to increase the airflow through nozzles 18 to a maximum value.
  • the supply of compressed air to the holes 14 is stopped and air is drawn from chambers 17 so that the primary and secondary airflows are increased.
  • the chambers 19 are pressurized to reduce the dilution airflow.
  • FIG. 1 shows airflow conditions obtaining at low throughput and the lower half shows conditions at high throughput.
  • the fluidic devices used in nozzles 15 and 18 may alternatively take the form of vortex amplifiers as shown in FIGS. 3 and 4 in which there are control drillings which can direct tangential jets of air into the nozzles to create swirl which will effectively reduce the area of the nozzles.
  • the sizes of the control drillings and the control pressures used would be chosen so that there is no change in the overall resistance of the combustion chamber to airflow when changeover from one flow condition to the other takes place.
  • a combustion chamber which includes a primary air inlet duct, secondary air inlet nozzles and dilution air inlet nozzles, each air entry point incorporating a fluidic airflow control device arranged so that the proportions of the total airflow through the respective air entry points can be varied without variation of the total resistance of the combustion chamber to airflow therethrough.
  • a combustion chamber as claimed in claim 1 in which the secondary air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings in the upstream sides thereof through which air may be injected or withdrawn so as respectively to reduce or increase the effective areas of the nozzles.
  • a combustion chamber as claimed in claim 1 in which the dilution air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings through which, in use, high-pressure air is injected tangentially into the nozzles.
  • a combustion chamber as claimed in claim 1 in which the dilution air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings in the upstream sides thereof through which air may be injected or withdrawn so as respectively to reduce or increase the effective areas of the nozzles.
  • a combustion chamber as claimed in claim 1 in which the secondary air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings through which in use high pressure air is injected tangentially through the nozzles.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A combustion chamber for a gas turbine engine has a primary air inlet and a number of additional air nozzles at intervals along the chamber. Each of the air inlets has one or more holes in the inlet wall, so that air may be injected into the inlets or extracted therefrom. Air injection effectively reduces the inlet areas and extraction increases the inlet areas. The airflow through each inlet, or group of inlets, may thus be controlled, the arrangement being such that, irrespective of the proportion of the total airflow which enters each inlet, the resistance to airflow through the combustion chamber does not vary. The holes in the inlet walls may be tangential to create a vortex within the inlet.

Description

United States Patent [72] lnventor Alan Joseph Gerrard Blackburn, England [21] AppLNo. 858,770
[22] Filed [45] Patented [73] Assignee Sept. 17, 1969 July 20, 1971 Joseph Lucas Industries Limited [54] COMBUSTION CHAMBERS FOR GAS TURBINE ENGINES 7 Claims, 4 Drawing Figs.
[52] US. Cl 60/39.65, 60/3923, 431/352 [51] lnt.Cl F02c9/14 [50] Field of Search 60/3965,
39.23, 39.29, 39.69; 137/13, 13.], 13.2, 81,5; 415/168,115,ll6,144,108;43l/350-353; 123/119, 119 C; 261/64, 69; 239/D1G. 3
[56] References Cited UNlTED STATES PATENTS 2,807,933 10/1957 Martin 60/3965 2,841,182 7/1958 Scala 60/3965 3,394,543 7/1968 Slattery 60/3965 FOREIGN PATENTS 738,006 10/1955 Great Britain 60/3965 Primary Examiner-Douglas Hart Attorney-Holman & Stern ABSTRACT: A combustion chamber for a gas turbine engine has a primary air inlet and a number of additional air'nozzles at intervals along the chamber. Each of the air inlets has one or more holes in the inlet wall, so that air may be injected into the inlets or extracted therefrom. Air injection effectively reduces the inlet areas and extraction increases the inlet areas. The airflow through each inlet, or group of inlets, may thus be controlled, the arrangement being such that, irrespective of the proportion of the total airflow which enters each inlet, the resistance to airflow through the combustion chamber does not vary. The holes in the inlet walls may be tangential to create a vortex within the inlet.
PATENTEH JUL P 0 mm SHEETIGFZ \A g l3 FIG. 2.
FICA.
INV NTOE 55 ,z fm
I -FTBRNEYS COMBUSTION CHAMBERS FOR GAS TURBINE ENGINES This invention relates to combustion chambers for gas turbine engines and hasas an object to provide a combustion chamber in a convenient form.
A combustion chamber in accordance with the invention has a primary air inlet duct, secondary air inlet nozzles and dilution air inlet nozzles, each air entry point incorporating a fluidic airflow control device arranged so that the proportions of the total airflow through the respective air entry points can be varied without variation of the total resistance of the combustion chamber to airflow therethrough.
In the accompanying drawings FIG. 1 is a fragmentary section through an annular combustion chamber incorporating an example of the invention,
FIG. 2 is a fragmentary enlargement of the part ringed in FIG. 1 and FIGS. 3 and 4 show, somewhat diagrammatically, two views of a part of an alternative embodiment.
The combustion chamber has an inner wall and an outer wall 11 shaped tojoin at a circular leading end in which primary air inlet ducts 12 are formed. Each primary air inlet duct has a swirler 13 which incorporates control jet holes 14 through which high-pressure air can be injected. These holes 14 are arranged around the outer periphery of the swirler 13 so that when compressed air is injected the effective area of the swirler is reduced fluidically to reduce the airflow through the primary air inlet ducts. Conversely, when air is bled off through the holes 14 the effective area of the swirler is increased fluidically to increase the airflow through the primary inlet duct. 7
The walls 10, ll of the combustion chamber adjacent the inlet end thereof are formed with secondary air inlet nozzles 15 constituted by inwardly directed flanges on the walls. In association with the nozzles 15 is an annular dished member 16 defining an annular air chamber 17. Drillings are formed in the flanges which open into the nozzles 15 on the upstream sides thereof. Thus when high-pressure air is applied to the chambers 17 jets will issue into the nozzles 15 and fluidically restrict flow of air into the combustion chamber. If on the other hand air is drawn from the chambers 17 the effective area of the nozzles 15 will be increased so that there will be increased airflow through the nozzles 15.
Downstream of the nozzles 15 the walls 10, 11 have dilution air inlet nozzles 18 similar to the nozzles 15. These nozzles 18 similarly have associated therewith fluidic flow control devices constituted by drillings in the nozzle flanges opening into chambers 19 to which air is supplied to decrease the effective areas of nozzles 18 and from which air is drawn to increase the effective area.
It will be noted that the provision of the fluidic control drillings on the upstream side only of the nozzles has the effect of changing the direction of the airflow through the nozzles as well as reducing the effective cross-sectional area of the nozzles.
At engine running conditions associated with the weaker overall air/fuel ratios for the combustion chamber, e.g., aircraft standoff flight conditions, the fluidic devices mentioned above are actuated to decrease flow through the primary and secondary air nozzles to a minimum value, whilst air is drawn from the chambers 19 to increase the airflow through nozzles 18 to a maximum value. At engine running conditions associated with the richer overall air/fuel ratios for the combustion chamber, e.g., aircraft takeoff conditions, the supply of compressed air to the holes 14 is stopped and air is drawn from chambers 17 so that the primary and secondary airflows are increased. The chambers 19 are pressurized to reduce the dilution airflow. This changes the pattern of airflow in the combustion chamber to increase the quantity of air available to the burners which would be situated in the centers of the respective swirlers 13. The change in direction of the secondary airflow causes increased reverse flow of secondary air to increase the rate ofintermixing of fuel and air. I
It IS to be noted that the upper half of FIG. 1 shows airflow conditions obtaining at low throughput and the lower half shows conditions at high throughput.
The fluidic devices used in nozzles 15 and 18 may alternatively take the form of vortex amplifiers as shown in FIGS. 3 and 4 in which there are control drillings which can direct tangential jets of air into the nozzles to create swirl which will effectively reduce the area of the nozzles.
ln either case the a eas of the nozzles, the sizes of the control drillings and the control pressures used would be chosen so that there is no change in the overall resistance of the combustion chamber to airflow when changeover from one flow condition to the other takes place.
Having thus described my invention what I claim as new and desire to secure by Letters Patent is:
l. A combustion chamber which includes a primary air inlet duct, secondary air inlet nozzles and dilution air inlet nozzles, each air entry point incorporating a fluidic airflow control device arranged so that the proportions of the total airflow through the respective air entry points can be varied without variation of the total resistance of the combustion chamber to airflow therethrough.
2. A combustion chamber as claimed in claim 1 in which the primary air inlet duct incorporates holes through which air may be injected or withdrawn so as respectively to increase or decrease the effective area of the duct.
3. A combustion chamber as claimed in claim 2 in which the said holes are formed in the outer periphery of the duct.
4. A combustion chamber as claimed in claim 1 in which the secondary air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings in the upstream sides thereof through which air may be injected or withdrawn so as respectively to reduce or increase the effective areas of the nozzles.
5. A combustion chamber as claimed in claim 1 in which the dilution air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings through which, in use, high-pressure air is injected tangentially into the nozzles.
6. A combustion chamber as claimed in claim 1 in which the dilution air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings in the upstream sides thereof through which air may be injected or withdrawn so as respectively to reduce or increase the effective areas of the nozzles.
7. A combustion chamber as claimed in claim 1 in which the secondary air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings through which in use high pressure air is injected tangentially through the nozzles.

Claims (7)

1. A combustion chamber which includes a primary air inlet duct, secondary air inlet nozzles and dilution air inlet nozzles, each air entry point incorporating a fluidic airflow control device arranged so that the proportions of the total airflow through the respective air entry points can be varied without variation of the total resistance of the combustion chamber to airflow therethrough.
2. A combustion chamber as claimed in claim 1 in which the primary air inlet duct incorporates holes through which air may be injected or withdrawn so as respectively to increase or decrease the effective area of the duct.
3. A combustion chamber as claimed in claim 2 in which the said holes are formed in the outer periphery of the duct.
4. A combustion chamber as claimed in claim 1 in which the secondary air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings in the upstream sides thereof through which air may be injected or withdrawn so as respectively to reduce or increase the effective areas of the nozzles.
5. A combustion chamber as claimed in claim 1 in which the dilution air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings through which, in use, high-pressure air is injected tangentially into the nozzles.
6. A combustion chamber as claimed in claim 1 in which the dilution air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings in the upstream sides thereof through which air may be injected or withdrawn so as respectively to reduce or increase the effective areas of the nozzles.
7. A combustion chamber as claimed in claim 1 in which the secondary air inlet nozzles are constituted by inwardly directed flanges on the walls of the chamber, said flanges being formed with drillings through which in use high pressure air is injected tangentially through the nozzles.
US858770A 1968-09-20 1969-09-17 Combustion chambers for gas turbine engines Expired - Lifetime US3593518A (en)

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Cited By (30)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3814575A (en) * 1973-04-25 1974-06-04 Us Air Force Combustion device
US3859786A (en) * 1972-05-25 1975-01-14 Ford Motor Co Combustor
US4021186A (en) * 1974-06-19 1977-05-03 Exxon Research And Engineering Company Method and apparatus for reducing NOx from furnaces
US4028044A (en) * 1974-10-07 1977-06-07 Rolls-Royce (1971) Limited Fuel burners
US4036582A (en) * 1974-11-02 1977-07-19 Motoren- Und Turbinen-Union Munchen Gmbh Combustion chamber for gas turbine power plants having devices for the gaseous processing of the fuel being introduced therein
US4052144A (en) * 1976-03-31 1977-10-04 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Fuel combustor
US4115050A (en) * 1975-10-09 1978-09-19 J. Eberspacher Burner construction and method for burning liquid and/or gaseous fuel
US4276018A (en) * 1979-05-30 1981-06-30 Davey Compressor Co. Mobile heater
US4311451A (en) * 1977-09-13 1982-01-19 Hitachi, Ltd. Burner
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5235805A (en) * 1991-03-20 1993-08-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Gas turbine engine combustion chamber with oxidizer intake flow control
US5289686A (en) * 1992-11-12 1994-03-01 General Motors Corporation Low nox gas turbine combustor liner with elliptical apertures for air swirling
US5322026A (en) * 1992-12-21 1994-06-21 Bay Il H Waste combustion chamber with tertiary burning zone
WO1999032828A1 (en) * 1997-12-18 1999-07-01 The Secretary Of State For Defence Fuel injector
US20060130486A1 (en) * 2004-12-17 2006-06-22 Danis Allen M Method and apparatus for assembling gas turbine engine combustors
US20090205309A1 (en) * 2006-08-30 2009-08-20 Deutsches Zentrum Fuer Luft- Und Raumfahrt E.V. Method for controlling the combustion in a combustion chamber and combustion chamber device
US20100218503A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US20100218504A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
US20130019604A1 (en) * 2011-07-21 2013-01-24 Cunha Frank J Multi-stage amplification vortex mixture for gas turbine engine combustor
CN103032891A (en) * 2013-01-04 2013-04-10 中国科学院工程热物理研究所 Multi-vortex combustion method
EP2236930A3 (en) * 2009-03-30 2013-07-31 United Technologies Corporation Combustor for gas turbine engine
US20140190171A1 (en) * 2013-01-10 2014-07-10 Honeywell International Inc. Combustors with hybrid walled liners
US20140373568A1 (en) * 2013-06-25 2014-12-25 Unique Gas Products Ltd. Direct venting system for free-standing propane powered absorption refrigerator
US20160123594A1 (en) * 2014-11-04 2016-05-05 United Technologies Corporation Low lump mass combustor wall with quench aperture(s)
US20160186998A1 (en) * 2013-08-30 2016-06-30 United Technologies Corporation Contoured dilution passages for gas turbine engine combustor
US20160201908A1 (en) * 2013-08-30 2016-07-14 United Technologies Corporation Vena contracta swirling dilution passages for gas turbine engine combustor
WO2017133819A1 (en) * 2016-02-01 2017-08-10 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber having a wall contour
US20180283695A1 (en) * 2017-04-03 2018-10-04 United Technologies Corporation Combustion panel grommet
US20220390115A1 (en) * 2021-06-07 2022-12-08 General Electric Company Combustor for a gas turbine engine
US20230296250A1 (en) * 2022-03-21 2023-09-21 General Electric Company Turbine engine combustor and combustor liner

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Publication number Priority date Publication date Assignee Title
GB738006A (en) * 1952-07-12 1955-10-05 Rolls Royce Improvements in or relating to gas turbine engines
US2807933A (en) * 1954-04-01 1957-10-01 Martin Peter Combustion chambers
US2841182A (en) * 1955-12-29 1958-07-01 Westinghouse Electric Corp Boundary layer fluid control apparatus
US3394543A (en) * 1966-04-29 1968-07-30 Rolls Royce Gas turbine engine with means to accumulate compressed air for auxiliary use

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB738006A (en) * 1952-07-12 1955-10-05 Rolls Royce Improvements in or relating to gas turbine engines
US2807933A (en) * 1954-04-01 1957-10-01 Martin Peter Combustion chambers
US2841182A (en) * 1955-12-29 1958-07-01 Westinghouse Electric Corp Boundary layer fluid control apparatus
US3394543A (en) * 1966-04-29 1968-07-30 Rolls Royce Gas turbine engine with means to accumulate compressed air for auxiliary use

Cited By (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3859786A (en) * 1972-05-25 1975-01-14 Ford Motor Co Combustor
US3814575A (en) * 1973-04-25 1974-06-04 Us Air Force Combustion device
US4021186A (en) * 1974-06-19 1977-05-03 Exxon Research And Engineering Company Method and apparatus for reducing NOx from furnaces
US4028044A (en) * 1974-10-07 1977-06-07 Rolls-Royce (1971) Limited Fuel burners
US4036582A (en) * 1974-11-02 1977-07-19 Motoren- Und Turbinen-Union Munchen Gmbh Combustion chamber for gas turbine power plants having devices for the gaseous processing of the fuel being introduced therein
US4115050A (en) * 1975-10-09 1978-09-19 J. Eberspacher Burner construction and method for burning liquid and/or gaseous fuel
US4052144A (en) * 1976-03-31 1977-10-04 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Fuel combustor
US4311451A (en) * 1977-09-13 1982-01-19 Hitachi, Ltd. Burner
US4276018A (en) * 1979-05-30 1981-06-30 Davey Compressor Co. Mobile heater
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5235805A (en) * 1991-03-20 1993-08-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Gas turbine engine combustion chamber with oxidizer intake flow control
US5289686A (en) * 1992-11-12 1994-03-01 General Motors Corporation Low nox gas turbine combustor liner with elliptical apertures for air swirling
US5322026A (en) * 1992-12-21 1994-06-21 Bay Il H Waste combustion chamber with tertiary burning zone
WO1999032828A1 (en) * 1997-12-18 1999-07-01 The Secretary Of State For Defence Fuel injector
US6474569B1 (en) 1997-12-18 2002-11-05 Quinetiq Limited Fuel injector
US20060130486A1 (en) * 2004-12-17 2006-06-22 Danis Allen M Method and apparatus for assembling gas turbine engine combustors
US20090205309A1 (en) * 2006-08-30 2009-08-20 Deutsches Zentrum Fuer Luft- Und Raumfahrt E.V. Method for controlling the combustion in a combustion chamber and combustion chamber device
US8171740B2 (en) 2009-02-27 2012-05-08 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
US8141365B2 (en) * 2009-02-27 2012-03-27 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US20100218503A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US20100218504A1 (en) * 2009-02-27 2010-09-02 Honeywell International Inc. Annular rich-quench-lean gas turbine combustors with plunged holes
EP2224170A3 (en) * 2009-02-27 2018-03-28 Honeywell International Inc. Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
EP2236930A3 (en) * 2009-03-30 2013-07-31 United Technologies Corporation Combustor for gas turbine engine
US9222674B2 (en) * 2011-07-21 2015-12-29 United Technologies Corporation Multi-stage amplification vortex mixture for gas turbine engine combustor
US20130019604A1 (en) * 2011-07-21 2013-01-24 Cunha Frank J Multi-stage amplification vortex mixture for gas turbine engine combustor
CN103032891B (en) * 2013-01-04 2015-07-29 中国科学院工程热物理研究所 A kind of many eddy combustion methods
CN103032891A (en) * 2013-01-04 2013-04-10 中国科学院工程热物理研究所 Multi-vortex combustion method
US20140190171A1 (en) * 2013-01-10 2014-07-10 Honeywell International Inc. Combustors with hybrid walled liners
US20140373568A1 (en) * 2013-06-25 2014-12-25 Unique Gas Products Ltd. Direct venting system for free-standing propane powered absorption refrigerator
US20160186998A1 (en) * 2013-08-30 2016-06-30 United Technologies Corporation Contoured dilution passages for gas turbine engine combustor
US20160201908A1 (en) * 2013-08-30 2016-07-14 United Technologies Corporation Vena contracta swirling dilution passages for gas turbine engine combustor
US11112115B2 (en) * 2013-08-30 2021-09-07 Raytheon Technologies Corporation Contoured dilution passages for gas turbine engine combustor
US20160123594A1 (en) * 2014-11-04 2016-05-05 United Technologies Corporation Low lump mass combustor wall with quench aperture(s)
US10451281B2 (en) * 2014-11-04 2019-10-22 United Technologies Corporation Low lump mass combustor wall with quench aperture(s)
US10670270B2 (en) 2016-02-01 2020-06-02 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with wall contouring
WO2017133819A1 (en) * 2016-02-01 2017-08-10 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber having a wall contour
US20180283695A1 (en) * 2017-04-03 2018-10-04 United Technologies Corporation Combustion panel grommet
US20220390115A1 (en) * 2021-06-07 2022-12-08 General Electric Company Combustor for a gas turbine engine
US11959643B2 (en) * 2021-06-07 2024-04-16 General Electric Company Combustor for a gas turbine engine
US20230296250A1 (en) * 2022-03-21 2023-09-21 General Electric Company Turbine engine combustor and combustor liner

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SE357028B (en) 1973-06-12
GB1278590A (en) 1972-06-21
DE1947762B2 (en) 1975-07-24
DE1947762A1 (en) 1970-04-02
FR2018565A1 (en) 1970-05-29

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