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Publication numberUS3593518 A
Publication typeGrant
Publication dateJul 20, 1971
Filing dateSep 17, 1969
Priority dateSep 20, 1968
Also published asDE1947762A1, DE1947762B2
Publication numberUS 3593518 A, US 3593518A, US-A-3593518, US3593518 A, US3593518A
InventorsGerrard Alan Joseph
Original AssigneeLucas Industries Ltd
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Combustion chambers for gas turbine engines
US 3593518 A
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Description  (OCR text may contain errors)

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.

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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.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2807933 *Mar 30, 1955Oct 1, 1957Peter MartinCombustion chambers
US2841182 *Dec 29, 1955Jul 1, 1958Westinghouse Electric CorpBoundary layer fluid control apparatus
US3394543 *Mar 30, 1967Jul 30, 1968Rolls RoyceGas turbine engine with means to accumulate compressed air for auxiliary use
GB738006A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3814575 *Apr 25, 1973Jun 4, 1974Us Air ForceCombustion device
US3859786 *May 25, 1972Jan 14, 1975Ford Motor CoCombustor
US4021186 *May 15, 1975May 3, 1977Exxon Research And Engineering CompanyMethod and apparatus for reducing NOx from furnaces
US4028044 *Sep 30, 1975Jun 7, 1977Rolls-Royce (1971) LimitedFuel burners
US4036582 *Oct 31, 1975Jul 19, 1977Motoren- Und Turbinen-Union Munchen GmbhCombustion chamber for gas turbine power plants having devices for the gaseous processing of the fuel being introduced therein
US4052144 *Mar 31, 1976Oct 4, 1977The United States Of America As Represented By The Administrator Of The National Aeronautics And Space AdministrationFuel combustor
US4115050 *Sep 13, 1976Sep 19, 1978J. EberspacherBurner construction and method for burning liquid and/or gaseous fuel
US4276018 *May 30, 1979Jun 30, 1981Davey Compressor Co.Mobile heater
US4311451 *Sep 13, 1978Jan 19, 1982Hitachi, Ltd.Burner
US5109671 *Dec 5, 1989May 5, 1992Allied-Signal Inc.Combustion apparatus and method for a turbine engine
US5235805 *Mar 12, 1992Aug 17, 1993Societe 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
US5289686 *Nov 12, 1992Mar 1, 1994General Motors CorporationLow nox gas turbine combustor liner with elliptical apertures for air swirling
US5322026 *Dec 21, 1992Jun 21, 1994Bay Il HWaste combustion chamber with tertiary burning zone
US6474569Dec 18, 1998Nov 5, 2002Quinetiq LimitedFuel injector
US8141365 *Feb 27, 2009Mar 27, 2012Honeywell International Inc.Plunged hole arrangement for annular rich-quench-lean gas turbine combustors
US8171740Feb 27, 2009May 8, 2012Honeywell International Inc.Annular rich-quench-lean gas turbine combustors with plunged holes
US20060130486 *Dec 17, 2004Jun 22, 2006Danis Allen MMethod and apparatus for assembling gas turbine engine combustors
US20130019604 *Jul 21, 2011Jan 24, 2013Cunha Frank JMulti-stage amplification vortex mixture for gas turbine engine combustor
EP2236930A2 *Mar 30, 2010Oct 6, 2010United Technologies CorporationCombustor for gas turbine engine
WO1999032828A1 *Dec 18, 1998Jul 1, 1999Secr DefenceFuel injector
Classifications
U.S. Classification60/752, 60/39.23, 431/352
International ClassificationF23R3/06, F23R3/02, F23R3/04, F23R3/26
Cooperative ClassificationF23R3/06, F23R3/04, F23R3/26
European ClassificationF23R3/06, F23R3/26, F23R3/04