Search Images Maps Play YouTube News Gmail Drive More »
Sign in
Screen reader users: click this link for accessible mode. Accessible mode has the same essential features but works better with your reader.

Patents

  1. Advanced Patent Search
Publication numberUS3099134 A
Publication typeGrant
Publication dateJul 30, 1963
Filing dateDec 27, 1960
Priority dateDec 24, 1959
Publication numberUS 3099134 A, US 3099134A, US-A-3099134, US3099134 A, US3099134A
InventorsCalder Peter Henry, Lancaster Derek Joh Stansfield
Original AssigneeHavilland Engine Co Ltd
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Combustion chambers
US 3099134 A
Abstract  available in
Images(2)
Previous page
Next page
Claims  available in
Description  (OCR text may contain errors)

y 30 1963 RH. CALDER Em 3,099,134

COMBUSTION CHAMBERS 2 Sheets$heet 1 Filed Dec. 2'7, 1960 \NVENTOQS PETER H. CALDER JQHN S. LANcasTER ATTORNEY July 30, 1963 P. H. CALDER ETAL 3,099,134

COMBUSTION CHAMBERS Filed Dec. 2'7, 1960 2 Sheets-Sheet 2 Fla l9 l6 /7c \NVENTORS PETER H. CALDER JOHN S. LANCASTER ATTORNEY United States Patent Filed Dec. 27, 1960, Ser. No. 78,352

Claims priority, application Great Britain Dec. 24, 1959 3 Claims. (Cl. 6039.65)

This invention relates to annular combustion chambers for use in gas turbines, for example for aircraft, industrial, or marine applications.

In conventional annular combustion chambers the gases pass through the combustion chamber travelling to the turbine nozzle ring in a substantially axial direction. This has various disadvantages and more particularly if the gas temperature is not uniform around the whole turbine nozzle ring, the nozzle ring or the turbine may be overheated at localised areas. The flow pattern in the combustion chamber may be very susceptible to changes in compressor outlet pressure distribution, and if any change in operating conditions occurs, the flame may go out or the radial temperature distribution may be such that the turbine rotor blades are severely overheated in highly stressed regions.

In all gas turbine combustion chambers the flame in the primary zone is stabilised by the recirculation of hot gases to give continuous piloting of incoming fuel/air mixture. In annular chambers this flow recirculation usually takes the form of a number of toroidal vortices spaced around the circumference of the annulus. Each toroid is usually maintained by combination of airflows from a swirler and from both the outer and the inner flame tubes. The symmetry and the balance of the flow in the toroid is dependent upon careful matching of swirler and flame tube a-irflows. Minor changes to swirler airflow distribution or to the pressure drops across outer and inner flame tubes can cause major distortion in the primary zone flow recirculation. Such distortions are caused by variations in compressor outlet flow distribution due to changes in engine operating conditions. Primary zone flow distortions cause drastic changes in combustion chamber outlet temperature distribution and can easily result in turbine blade failure due to overheating at critically-stressed sections.

The principal object of the present invention is to devise a primary zone flow recirculation which is less sensitive to the effects of compressor outlet flow distribution changes and which is basically stable.

Now according to the present invention, an annular combustion chamber for a gas turbine engine, includes air inlet apertures spaced around radially inner and/or outer walls defining an annular combustion space, as distinct from the end walls, and arranged to impart a velocity to the gases within the combustion space in a direction having a substantial circumferential component around the annular combustion chamber.

The bodily rotation of the combustion gases within the combustion chamber provides several advantages. In the first place the rotation of the air will assist in obtaining a good mixing and distribution of the fuel-air mixture in the combustion zone and this will result in a stable flame, secondly the rotation will give good mixing 3,099,134 Patented July 30, 1963 "ice of the combustion gases and will tend to give a uniform gas exit temperature. Thirdly the rotational movement of the air and of the combustion gases will result in a flow path of increased length which in turn will result in more complete combustion of the fuel before the gases reach the turbine. In some cases it may be possible in some instances to shorten the length of the combustion unit.

The flow distribution may be basically stable by virtue of the combined effect of the rotation of the air in the primary zone circumferentially, and a subsid-ary vortex flow of air through the inlet aperture towards the up stream end of the combustion space due to the low pressure there, and thence out of the combustion space along the other side of the flame tube.

Conveniently, at least certain of the inlet apertures extend with axes at a substantial angle to the radius of the combustion space in order to impart the circumferential velocity.

It is preferred that substantially all the air enters the combustion chamber through the said air inlet apertures in the inner and outer Walls. A certain amount of air will be used conventionally to cool the fuel nozzles and the combustion space walls.

It is also preferred that there is no swirler through which air enters the combustion space. This means that the flow path in the primary zone is substantially dependent on air directed into the primary zone from the flame tube walls. Changes in mass flow distribution between outer and inner annular passages supplying air respectively to the inner and outer flame tube may then cause the velocity and mixture strength in the primary zone to change without altering the =flow path.

The inlet apertures may be only in the inner wall, or only in the outer wall defining the combustion space, in order that the air from the compressor does not have to be divided into two or more flow paths before entering the combustion space. On the other hand, in some cases, stability can be achieved by having air inlet apertures in both the inner and outer walls, of which at least some of the apertures in the inner wall or the outer wall only are arranged to direct the air radially, rather than with a circumferential component.

In one form of the invention, the combustion chamber comprises two co-axial annular combustion spaces radially spaced apart th an air supply duct between them from which air enters both the combustion spaces through inlet apertures in the inner wall of the outer combustion space, and the outer wall of the inner combustion space.

It is preferred that the fuel, or some of the fuel is supplied to the combustion space through nozzles in the end wall, but it is also possible to supply some or all of the fuel through other nozzles, for example nozzles which lie within the air inlet apertures, and which may be directed radially, or with a circumferential component.

The invention may be carried into practice in various ways, and certain embodiments will now be described by way of example as applied to an annular combustion chamber arranged to be connected between a compressor and a turbine to receive air from the compressor and to deliver to the nozzle ring of the turbine the products of combustion.

Reference will be made to the accompanying drawings, in which:

FlGU-RE l is a diagrammatic section of a part of one embodiment of combustion chamber,

FIGURE 2 is a transverse section on the line -IIII in FIGURE 1,

FIGURE 3 is a diagrammatic section corresponding to FIGURE 11, of a second embodiment of combustion chamber in which the air enters only through the outer wall, and

FIGURE 4- is a section corresponding to FIGURES 1 and 3 of a third embodiment of combustion chamber having two radially-spaced concentric annular combustion spaces.

The combustion chamber comprises inner and outer generally cylindrical pressure resisting casings 11 and 12, defining an annular chamber, together with an annular flame tube within this chamber and spaced from the walls 11 and 12. The flame tube comprises outer and inner generally cylindrical, or part-conical walls 13 and '14, and an end wall 15 at the upstream end, which wall is of semi-circular profile as seen in FIGURE 1, and is connected to the upstream ends of the inner and outer walls.

Air enters the annular chamber 11, 12 from the compressor as shown by the arrow 16, and divides into two streams radially outside and inside the annular flame tube 13, 14, 15. The air enters the combustion space in the flame tube through air inlet apertures 17 in the outer and inner sidewalls 13 and 14. The air entering through the upstream apertures 1712 may be considered to be primary air. Air entering through apertures 17b further downstream may be considered to be secondary air, and air entering through down stream apertures 17c dilution air, although there is no strict distinction between primary and secondary air, and secondary and dilution air. The

primary air through the upstream apertures 17 is responsible for maintaining the flame stable so that the incoming fuel is effectively piloted.

The apertures l7 comprise short open pipes or ducts, as shown in FIGURE 2. These all have a tangential or circumferential component so that the chage will tend to be bodily rotated circumferentially around the combustion space. There may also be even a slight axial component upstream or downstream if it is found to be necessary.

The general flow of the combustion gases around the flame tube and a continuous toroidal vortex which exists at the upstream end of the combustion space, have been found to give an air flow pattern in the combustion space which is extremely stable in spite of changes in conditions, and in particular in spite of variations in the compressor outlet pressure of flow distribution, so that the effective piloting of the incoming fuel, and a stable temperature pattern along the turbine blades can be achieved.

It was suggested above that all the apertures should be directed with a tangential component as shown in FIG- URE 2, but in fact, in some cases, it may be that some apertures in the downstream rows 17b and 170 may be directed radially, and in some circumstances a more stable flow pattern still may be obtained if the rotation of the charge is achieved only by inclined apertures in the outer wall or in the inner wall, while the opposite apertures are radial. (This is shown dotted in FIGURE 2.)

In the embodiment shown in FIGURES 3, where similar parts have the same reference numerals as in FIGURES 1 and 2, the air inlet apertures through the wall of the flame tube lead only through the outer wall 13, and almost all of the air from the compressor passes into the annular duct between this wall 13, and the outer pressure resistant wall 11 of the combustion chamber. The upstream rows of air inlet apertures 17a are inclined tangentially in a manner similar to that shown in FIG- URE 2, While the downstream rows are directed radially. Once again, the fuel is supplied through a ring of nozzles such as 19 arranged in the end wall of the flame tube from an annular fuel gallery 21 surrounding the combustion chamber.

In the embodiment of FIGURE 4 the flame tube is formed to define two co-axial radially-spaced combustion spaces, 22 and 23, leading to a common exhaust manifold 24, and the air from the compressor at 16 is led to a central annular duct 25 radially between the spaces 22 and 23. Air enters the outer combustion space 23 through the inner wall only, and enters the inner combustion space 23 through the outer wall only, and once again the apertures 17a in the upstream rows are inclined with a circumferential component. Fuel from a common supply gallery 21 enters the flame tubes through rings of fuel nozzles 19 in the respective end walls.

The embodiments of FIGURES 3 and 4 have the advantage that the air from the compressor does not have to be split into two or three ducts with bends around which energy is lost, but is fed principally into a single straight duct only. This gives better diffusion of the air.

It has been said that the air enters the combustion space only through the side walls, but it is pointed out that in making this remark no notice has been taken of small quantities of air supplied primarily for cooling the fuel nozzles or the flame tube walls through small apertures in the conventional manner.

It is also possible that fuel may be supplied in other ways, for example through radially-directed or circumferentially inclined nozzles leading through certain of the air apertures 17, but it is preferred that the fuel enters through the end wall as shown.

By supplying the air only through the side walls, and with a circumferential component and through no swirler, it has been found that very stable flame patterns can be achieved in spite of considerable variations in compressor characteristics.

What we claim as our invention and desire to secure by Letters Patent is:

1. An annular combustion chamber for a gas turbine engine, including air inlet apertures spaced around both the radially inner and radially outer walls defining an annular com-bustion space, as distinct from the end walls, the air inlet apertures comprising a series of primary air apertures at the upstream end of the combustion chamber and a series of secondary air apertures further downstream of the primary air apertures, the primary air apertures in one of said walls being directed in a plane transverse to the axis of the combustion chamber, but with a substantial circumferential component of direction to establish a rotating burning charge and the primary air apertures in the other of said walls being arranged to direct the air radially rather than with a circumferential component, and fuel nozzles for directing fuel directly into the rotating burning charge in the combustion space.

2. A combustion chamber as claimed in claim 1 comprising two co-axial combustion spaces radially spaced apart with an air supply duct between them from which substantially all the air enters the combustion spaces through inlet apertures in the inner wall of the outer combustion space, and the outer wall of the inner combustion space.

3. An annular combustion chamber in combination with means defining an annular discharge path from an air compressor and walls defining an annular region connected through the annular discharge path, and embracing the annular combustion chamber, including air inlet apertures spaced around at least one of the radially inner and radially outer walls defining an annular combustion space, as distinct from the end Walls, the air inlet aperture comprising a series of primary air apertures at the upstream end of the combustion chamber, which apertures are directed substantially in a plane transverse to the axis of the combustion chamber, but with a substantial circumferential component of direction to establish a recirculating annular-1y rotating burning charge and a series of secondary air apertures further downstream of the primary air apertures, and fuel nozz1es for directing fuel dircctly into the rotating burning charge in the cornbustion space.

References Cited in the fi1e of this patent UNETED STATES PATENTS 2,576,046 Scarth Nov. 20, 1951 2,579,614 Ray Dec. 25, 1951 2,588,728 Hundstad Mar. 11, 1952 6 Stalker Feb. 10, 1953 Nathan May 12, 1953 Krejei Nov. 17, 1953 Bloomer Sept. 27, 1955 Johnson Mar. 29, 1960 Benson Sept. 6, 1960 FORETGN PATENTS France Oct. 21, 1920

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2576046 *Apr 29, 1946Nov 20, 1951Phillips Petroleum CoDouble-walled annular combustion chamber with turbine shaft air jacket
US2579614 *Jun 23, 1944Dec 25, 1951Allis Chalmers Mfg CoCombustion chamber with rotating fuel and air stream surrounding a flame core
US2588728 *Jun 14, 1948Mar 11, 1952Us NavyCombustion chamber with diverse combustion and diluent air paths
US2627719 *Jun 13, 1947Feb 10, 1953Edward A StalkerGas turbine combustion chamber having controlled laminar flow of air for combustion and insulation
US2637974 *Sep 5, 1945May 12, 1953Power Jets Res & Dev LtdCombustion apparatus for an air stream and propulsive system
US2659201 *Nov 26, 1947Nov 17, 1953Phillips Petroleum CoGas turbine combustion chamber with provision for turbulent mixing of air and fuel
US2718757 *Jan 17, 1951Sep 27, 1955Lummus CoAircraft gas turbine and jet
US2930192 *Dec 7, 1953Mar 29, 1960Gen ElectricReverse vortex combustion chamber
US2951339 *Mar 31, 1959Sep 6, 1960United Aircraft CorpCombustion chamber swirler
FR512723A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3643430 *Mar 4, 1970Feb 22, 1972United Aircraft CorpSmoke reduction combustion chamber
US3874169 *Sep 27, 1973Apr 1, 1975Stal Laval Turbin AbCombustion chamber for gas turbines
US3968644 *Aug 19, 1974Jul 13, 1976Motoren- Und Turbinen-Union Munchen GmbhFuel admitting and conditioning means on combustion chambers for gas turbine engines
US3974647 *Aug 26, 1974Aug 17, 1976United Technologies CorporationCombustion instability reduction device having swirling flow
US4008977 *Sep 19, 1975Feb 22, 1977United Technologies CorporationCompressor bleed system
US4016718 *Jul 21, 1975Apr 12, 1977United Technologies CorporationGas turbine engine having an improved transition duct support
US4054028 *Aug 28, 1975Oct 18, 1977Mitsubishi Jukogyo Kabushiki KaishaFuel combustion apparatus
US4087963 *Apr 23, 1976May 9, 1978Phillips Petroleum CompanyCombustor for low-level NOx and CO emissions
US4201047 *Apr 4, 1978May 6, 1980Morgan J RandolphLow emission combustors
US4301657 *May 3, 1979Nov 24, 1981Caterpillar Tractor Co.Gas turbine combustion chamber
US4590769 *Jan 12, 1981May 27, 1986United Technologies CorporationHigh-performance burner construction
US4621499 *Jul 11, 1985Nov 11, 1986Hitachi, Ltd.Gas turbine combustor
US4928479 *Dec 28, 1987May 29, 1990Sundstrand CorporationAnnular combustor with tangential cooling air injection
US5070700 *Mar 5, 1990Dec 10, 1991Rolf Jan MowillLow emissions gas turbine combustor
US5085039 *Dec 7, 1989Feb 4, 1992Sundstrand CorporationCoanda phenomena combustor for a turbine engine
US5140807 *Feb 1, 1991Aug 25, 1992Sundstrand CorporationAir blast tube impingement fuel injector for a gas turbine engine
US5150570 *Dec 21, 1989Sep 29, 1992Sundstrand CorporationUnitized fuel manifold and injector for a turbine engine
US5163284 *Feb 7, 1991Nov 17, 1992Sundstrand CorporationDual zone combustor fuel injection
US5174108 *Sep 23, 1991Dec 29, 1992Sundstrand CorporationTurbine engine combustor without air film cooling
US5177955 *Feb 7, 1991Jan 12, 1993Sundstrand Corp.Dual zone single manifold fuel injection system
US5177956 *Feb 6, 1991Jan 12, 1993Sundstrand CorporationUltra high altitude starting compact combustor
US5205117 *Feb 1, 1991Apr 27, 1993Sundstrand CorporationHigh altitude starting two-stage fuel injection
US5220794 *Jun 22, 1990Jun 22, 1993Sundstrand CorporationImproved fuel injector for a gas 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
US5261224 *Nov 2, 1992Nov 16, 1993Sundstrand CorporationHigh altitude starting two-stage fuel injection apparatus
US5263316 *Dec 21, 1989Nov 23, 1993Sundstrand CorporationTurbine engine with airblast injection
US5277022 *Oct 23, 1992Jan 11, 1994Sundstrand CorporationAir blast fuel injecton system
US5317864 *Sep 30, 1992Jun 7, 1994Sundstrand CorporationTangentially directed air assisted fuel injection and small annular combustors for turbines
US5377483 *Jan 7, 1994Jan 3, 1995Mowill; R. JanProcess for single stage premixed constant fuel/air ratio combustion
US5477671 *Jun 3, 1994Dec 26, 1995Mowill; R. JanSingle stage premixed constant fuel/air ratio combustor
US5479781 *Mar 7, 1995Jan 2, 1996General Electric CompanyLow emission combustor having tangential lean direct injection
US5481866 *Jun 14, 1994Jan 9, 1996Mowill; R. JanSingle stage premixed constant fuel/air ratio combustor
US5488829 *May 25, 1994Feb 6, 1996Westinghouse Electric CorporationMethod and apparatus for reducing noise generated by combustion
US5572862 *Nov 29, 1994Nov 12, 1996Mowill Rolf JanConvectively cooled, single stage, fully premixed fuel/air combustor for gas turbine engine modules
US5613357 *May 29, 1996Mar 25, 1997Mowill; R. JanStar-shaped single stage low emission combustor system
US5628182 *May 23, 1995May 13, 1997Mowill; R. JanStar combustor with dilution ports in can portions
US5638674 *Jul 5, 1994Jun 17, 1997Mowill; R. JanConvectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission
US5680765 *Jan 5, 1996Oct 28, 1997Choi; Kyung J.Lean direct wall fuel injection method and devices
US5765363 *Jan 6, 1997Jun 16, 1998Mowill; R. JanConvectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission
US5924276 *Jul 15, 1997Jul 20, 1999Mowill; R. JanPremixer with dilution air bypass valve assembly
US5987889 *Oct 9, 1997Nov 23, 1999United Technologies CorporationFuel injector for producing outer shear layer flame for combustion
US6220034Mar 3, 1998Apr 24, 2001R. Jan MowillConvectively cooled, single stage, fully premixed controllable fuel/air combustor
US6675587 *Mar 21, 2002Jan 13, 2004United Technologies CorporationCounter swirl annular combustor
US6925809Dec 14, 2001Aug 9, 2005R. Jan MowillGas turbine engine fuel/air premixers with variable geometry exit and method for controlling exit velocities
US7665309Sep 15, 2008Feb 23, 2010Siemens Energy, Inc.Secondary fuel delivery system
US7798765Apr 12, 2007Sep 21, 2010United Technologies CorporationOut-flow margin protection for a gas turbine engine
US8387398Aug 20, 2008Mar 5, 2013Siemens Energy, Inc.Apparatus and method for controlling the secondary injection of fuel
US8397514 *May 24, 2011Mar 19, 2013General Electric CompanySystem and method for flow control in gas turbine engine
US8601820Jun 6, 2011Dec 10, 2013General Electric CompanyIntegrated late lean injection on a combustion liner and late lean injection sleeve assembly
US8919127May 24, 2011Dec 30, 2014General Electric CompanySystem and method for flow control in gas turbine engine
US8919137Aug 5, 2011Dec 30, 2014General Electric CompanyAssemblies and apparatus related to integrating late lean injection into combustion turbine engines
US8925326May 24, 2011Jan 6, 2015General Electric CompanySystem and method for turbine combustor mounting assembly
US9010120Aug 5, 2011Apr 21, 2015General Electric CompanyAssemblies and apparatus related to integrating late lean injection into combustion turbine engines
US9140455Jan 4, 2012Sep 22, 2015General Electric CompanyFlowsleeve of a turbomachine component
US20080253884 *Apr 12, 2007Oct 16, 2008United Technologies CorporationOut-flow margin protection for a gas turbine engine
US20090084082 *Aug 20, 2008Apr 2, 2009Siemens Power Generation, Inc.Apparatus and Method for Controlling the Secondary Injection of Fuel
US20120297786 *Nov 29, 2012General Electric CompanySystem and method for flow control in gas turbine engine
USRE30160 *Jun 5, 1978Nov 27, 1979United Technologies CorporationSmoke reduction combustion chamber
USRE34962 *May 29, 1992Jun 13, 1995Sundstrand CorporationAnnular combustor with tangential cooling air injection
Classifications
U.S. Classification60/752, 60/746, 60/747, 60/758
International ClassificationF23R3/34, F23R3/04
Cooperative ClassificationY02T50/675, F23R3/34, F23R3/04
European ClassificationF23R3/34, F23R3/04