US3593518A - Combustion chambers for gas turbine engines - Google Patents
Combustion chambers for gas turbine engines Download PDFInfo
- 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
- Authority
- US
- United States
- Prior art keywords
- nozzles
- air
- combustion chamber
- air inlet
- inlet
- 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.)
- Expired - Lifetime
Links
Images
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
- F23R3/02—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.
Landscapes
- 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.
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB44805/68A GB1278590A (en) | 1968-09-20 | 1968-09-20 | Combustion chambers for gas turbine engines |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| US3593518A true US3593518A (en) | 1971-07-20 |
Family
ID=10434833
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| US858770A Expired - Lifetime US3593518A (en) | 1968-09-20 | 1969-09-17 | Combustion chambers for gas turbine engines |
Country Status (4)
| Country | Link |
|---|---|
| US (1) | US3593518A (en) |
| FR (1) | FR2018565A1 (en) |
| GB (1) | GB1278590A (en) |
| SE (1) | SE357028B (en) |
Cited By (31)
| 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 |
| US20100218504A1 (en) * | 2009-02-27 | 2010-09-02 | Honeywell International Inc. | Annular rich-quench-lean gas turbine combustors with plunged holes |
| US20100218503A1 (en) * | 2009-02-27 | 2010-09-02 | Honeywell International Inc. | Plunged hole arrangement for annular rich-quench-lean gas turbine combustors |
| 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 |
| EP4202301A1 (en) * | 2021-12-21 | 2023-06-28 | General Electric Company | Combustor with dilution openings |
| US20230296250A1 (en) * | 2022-03-21 | 2023-09-21 | General Electric Company | Turbine engine combustor and combustor liner |
Citations (4)
| 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 |
-
1968
- 1968-09-20 GB GB44805/68A patent/GB1278590A/en not_active Expired
-
1969
- 1969-09-17 US US858770A patent/US3593518A/en not_active Expired - Lifetime
- 1969-09-19 SE SE12892/69A patent/SE357028B/xx unknown
- 1969-09-22 FR FR6932136A patent/FR2018565A1/fr not_active Withdrawn
Patent Citations (4)
| 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 (42)
| 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 |
| US20100218503A1 (en) * | 2009-02-27 | 2010-09-02 | Honeywell International Inc. | Plunged hole arrangement for annular rich-quench-lean gas turbine combustors |
| US8141365B2 (en) * | 2009-02-27 | 2012-03-27 | 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 |
| US20130019604A1 (en) * | 2011-07-21 | 2013-01-24 | Cunha Frank J | Multi-stage amplification vortex mixture for gas turbine engine combustor |
| US9222674B2 (en) * | 2011-07-21 | 2015-12-29 | United Technologies Corporation | 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 |
| 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 |
| 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 |
| US20160201908A1 (en) * | 2013-08-30 | 2016-07-14 | United Technologies Corporation | Vena contracta swirling dilution passages for gas turbine engine combustor |
| US20160186998A1 (en) * | 2013-08-30 | 2016-06-30 | United Technologies Corporation | Contoured 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) |
| WO2017133819A1 (en) * | 2016-02-01 | 2017-08-10 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine combustion chamber having a wall contour |
| US10670270B2 (en) | 2016-02-01 | 2020-06-02 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine combustion chamber with wall contouring |
| 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 |
| EP4202301A1 (en) * | 2021-12-21 | 2023-06-28 | General Electric Company | Combustor with dilution openings |
| US20230296250A1 (en) * | 2022-03-21 | 2023-09-21 | General Electric Company | Turbine engine combustor and combustor liner |
| US12222104B2 (en) * | 2022-03-21 | 2025-02-11 | General Electric Company | Turbine engine combustor and combustor liner |
Also Published As
| Publication number | Publication date |
|---|---|
| FR2018565A1 (en) | 1970-05-29 |
| DE1947762A1 (en) | 1970-04-02 |
| GB1278590A (en) | 1972-06-21 |
| SE357028B (en) | 1973-06-12 |
| DE1947762B2 (en) | 1975-07-24 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US3593518A (en) | Combustion chambers for gas turbine engines | |
| US6474569B1 (en) | Fuel injector | |
| US3374624A (en) | Gas turbine engine combustion equipment | |
| US3285007A (en) | Fuel injector for a gas turbine engine | |
| US3299632A (en) | Combustion chamber for a gas turbine engine | |
| US3605939A (en) | Device for reducing the noise produced by fluid flow escaping from a nozzle | |
| KR950001074A (en) | Gas turbine burner | |
| US2979899A (en) | Flame spreading device for combustion equipments | |
| US6216466B1 (en) | Fuel-injection arrangement for a gas turbine combustor | |
| US4590769A (en) | High-performance burner construction | |
| US2990905A (en) | Jet noise suppression means | |
| US5097666A (en) | Combustor fuel injection system | |
| US3643431A (en) | Flow control devices | |
| EP0722065B1 (en) | Fuel injector arrangement for gas-or liquid-fuelled turbine | |
| GB780493A (en) | Improvements relating to combustion equipment for gas-turbine engines | |
| US2977760A (en) | Annular combustion chambers for use with compressors capable of discharging combustion supporting medium with a rotary swirl through an annular outlet | |
| GB1421399A (en) | Fuel injectors | |
| US2959003A (en) | Fuel burner | |
| US3483701A (en) | Gas turbine combustion equipment | |
| US3530667A (en) | Fuel injector for gas turbine engines | |
| US5339622A (en) | Gas turbine engine with improved water ingestion prevention | |
| GB1357533A (en) | Combustion equipment for gas turbine engines | |
| US3151453A (en) | Reheat combustion apparatus for a gas turbine engine | |
| US2851859A (en) | Improvements in combustion chambers for turbo-jet, turbo-prop and similar engines | |
| US5195315A (en) | Double dome combustor with counter rotating toroidal vortices and dual radial fuel injection |