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US2687010A - Combustion apparatus - Google Patents

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Publication number
US2687010A
US2687010A US5396748A US2687010A US 2687010 A US2687010 A US 2687010A US 5396748 A US5396748 A US 5396748A US 2687010 A US2687010 A US 2687010A
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Prior art keywords
air
combustion
fuel
chamber
apparatus
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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
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Ellis Hugh Stanley
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Power Jets Res and Dev Ltd
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Power Jets Res and Dev Ltd
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23CCOMBUSTION APPARATUS USING FLUENT FUEL
    • F23C99/00Subject-matter not provided for in other groups of this subclass
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23DBURNERS
    • F23D11/00Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04Air inlet arrangements
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23CCOMBUSTION APPARATUS USING FLUENT FUEL
    • F23C2700/00Special arrangements for combustion apparatus using fluent fuel
    • F23C2700/02Combustion apparatus using liquid fuel
    • F23C2700/023Combustion apparatus using liquid fuel without pre-vaporising means

Description

Aljg. 24, 1954 5. ELLls 2,687,010

. COMBUSTION APPARATUS 7 Filed Oct. 11, 1943 v s Sheeis-Sheet 1,

. Inventor; v I

. 1 1, a 2d, Attorneys H. s. ELLIS COMBUSTION APPARATUS Aug. 24, 1954 '3 Sheets-Sheet 2 Filed Oct. 11. 1948 Qua Invenfo l; WM

Attorney;

Aug. 24, 1954 H. s. ELLIS 3 2,537,010 COMBUSTION APPARATUS FiledOct. 11, 1948 :i Sheets-Sheet s l6 8 Bk H M 1 y l I- FIG4 , z zilngeggf I zmz Patented Aug. 24, 1954 COMBUSTION APPARATUS Hugh Stanley Ellis, East signor to Power Jets ment) company Grinstead, England, as-

(Research and Develop- Limited, London, England, a British Application October 11, 1948, Serial No. 53,967

Claims priority, application Great Britain November 3, 1947 2 Claims. (Cl. 60-39-65) This invention relates to combustion apparatus of the kind in which fuel is burnt in a continuous airstream. Whilst as will be seen after consideration of its details, the combustion apparatus of the invention has possible application in a wider field, the invention is primarily concerned, and is at present conceived to have A its maximum utility in connection with combustion apparatus in which special problems arise due to the necessity for supporting combustion by means of a fast moving air current involving a large mass flow, as for example, in gas turbine or/and jet propulsion power units, the descriptions fast moving being used here to indicate that the mean speed of the combustion-supporting air current in its general direction of flow past a combustion zone, calculated from the ratio air volume passing in unit time/cross sectional area of flow path, is substantially higher than the speed of flow propagation in the fuel/air mixture concerned. For hydrocarbon fuels burning in air the speedof flame propagation is considered as being of the order of one foot per second at atmospheric temperature; the invention, on the other hand, is especially applicable to combustion apparatus for gas turbine or/and jet propulsion units in which the speed of the air current in its general direction of flow past a combustion zone, calculated on the basis indicated, might be of an order as low as or as high as 300 feet per second or even more, depending on the design.

In certain applications of combustion apparatus in which the reduction of bulk and weight to a minimum are important considerations, it is of value to be able to keep the flame length as short as possible; this is notably the case in a gas turbine used as an aircraft power plant, in which the combustion apparatus is usually arranged annularly about the longitudinal axis of the plant, so that a small increase in the axial length of the combustion apparatus results in a disproportionate increase in the structural mass. A further, consideration in combustion apparatus of the kind indicated, which again is of particular importance in relation to gas turbines, is the attainment of an even air/fuel flow pattern, which ideally should be such that the distribution of fuel and air, and the combustion temperature, are uniform at all points in the cross sectional plane of the outlet from the apparatus. In addition, in the case of a gas turbine the combustion apparatus should be capable of satisfactory operation over a wide range of air mass flows and fuel ratios without the flame being extinuished.

With a view to enabling the satisfaction of the requirements mentioned in the foregoing the invention proposes a combustion apparatus in which the fuel to be burnt (whether liquid, gaseous or pulverulent) is injected in or near a region of the combustion-supporting air stress in which a component or reverse flow is afforded by creating in the air stream a whirl having the general form of a toroidal annulus lying in a plane transverse to the stream (that is to say, of an annular envelope which defines a more or less closed figure in diametrical cross section).

It is intended that the fuel should be injected in such a way as to secure as far as possible a uniform distribution of fuel over the whirling air flow, but the manner in which the injection will take place will depend upon the circumstances of the case. Thus, the fuel may be injected into the toroidal air whirl in a generally circumferential direction, that is to say, in a general sense tangentially to the annulus formed by the toroid, and at a plurality of points distributed around it. Alternatively the fuel may be injected at one or more points in a downstream or an upstream direction and more or less along or parallel with the axis about which the annulus of the toroid is formed, or radially inwardly or outwardly with respect to said annulus. In any of these arrangements, by providing for an appropriate location and divergence of the jet from a single injection nozzle, or an appropriately close spacing and arrangement of a plurality of fuel jets, a more or less uniform toroidal annulus of fuel/ air mixture may be produced, with fuel/air mixture leaving the whirl in the general direction of the air flow.

The number of fuel jets required to ensure reasonably uniform distribution of the fuel, and the location thereof, will be influenced by the form of the combustion chamber. In this connection the invention has particular application to the case of a combustion apparatus, such as is employed in some types of gas turbine power plant, in which there in a single annular combustion chamber receiving the whole of the output from a compressor and constituting the sole source of supply for a turbine but it can also be applied to the corresponding case in which the combustion apparatus is formed by a simple tubular combustion chamber or is sub-divided into a number of such combustion chambers operating in parallel. In either case a whirl having the general form of toroidal annulus may be produced in the air fiow by the provision of air inlets to the fuel injection zone suitably disposed to direct a series of air jets along mutually inshould take place at several circumferentially.

spaced points and in these circumstances the fuel jets may with advantage be arranged with their axes of injection in the circumferential direction and contained within the envelope of the toroidal air whirl, although desirably not coinciding with the centre of the whirl. Such an arrangement allows fuel. to be well distributed about the annulus with the use of a minimum number of fuel jets, but where uniformity of the fuel distribution is of less importance or a larger number of jets is permissible the injection may take place in a direction upstream, downstream, or radially inwardly or outwardly against the toroidal whirl.

the method of injection may be similarly to those described, the annulus formed by the toroidal whirl may be sufficiently small to allow the fuel to be distributed uniformly to it by a single jet directed upstream or downstream along the axis about which the toroid is formed.

The air introduced to form the toroidal whirl could constitute the whole of the primary air supply for supporting the initial stages of combustion, thus eliminating the need for a separate primary air supply. In any event it will be necestional broadly novel feature, for a combustion sary to introduce further or secondary air for the completion of combustion into the flame zone at successive points in the downstream direction. In an annular combustion chamber such introduction is commonly effected through both the outer and inner walls, and adequate air supply thereby-maintained; normally, however, a tubular combustion chamber permits of such introduction only through its outer wall and the core of the flame may, in consequence, be inadequately supplied with air.

The invention therefore provides, as an additionally broadly novel feature, for a combustion apparatus comprising a simple tubular combustion chamber, or group of such chambers operating in parallel, in which secondary air for the completion of combustion is fed, in each chamber, to an inner zone thereof extending for a substantial distance downstream from the region at which fuel is injected, so to maintain a supply of air to the core of the flame in the chamber. Thus, in a simple tubular combustion chamber in which a toroidal annulus of fuel/air mixture is created, air may be supplied through the centre of the annulus of the toroid to the core of a flame extending and converging downstream thereof. This may be effected in a tubular chamber having annular-1y disposed fuel injectors by providing at the upstream end of the chamber an axially extending inner tube having communication at its upstream end with the air supply and extending downstream as far as required for the supply of secondary combustion air.

In order that the various aspects of the invention may be more readily understood, they will now be described with reference to the accompanying drawings in which On the other hand, in the case of a single tubular combustion chamber, whilst Figure 1 represents diagrammatically a crosssection in a plane extending in the direction of the combustion-supporting air stream of the fuel injection zone of the combustion chamber of a combustion apparatus according to the inven tion.

Figure 2 is a cross section in a plane transverse to the combustion supporting airstream of the combustion chamber of Figure 1.

Figure 3 is a half axial cross-section of an annular combustion apparatus (that is having an annular outlet) according to the invention.

Figure 4 is an axial cross-section of a tubular combustion apparatus (that is having a non-annular outlet) according to the invention.

In Figure 1 the fuel-injection zone I is shielded from the main air stream (the direction of which is indicated by the arrow A) by the three walls, 2, 3, and s respectively, of a casing. Three groups of air inlets 5, 6 and l are provided in the walls 2, 3, and 4 respectively, situated in the manner shown. Air jets entering the respec tive groups of inlets are so directed as to act in conjunction to produce the required whirling air flow path as indicated generally by the arrows It is apparent that a combustion chamber whose upstream end comprises such a casing with the three walls 2, 3, 4 extending circumferentially to form a complete chamber annulus about an axis parallel with the air stream (as shown partly in elevation in Figure 2) and with the groups of air inlets extending around its en'- tire circumferences, will have a toroidal air whirl formed within it. Arrangements of air inlets other than that shown could be used to produce a similar effect. The fuel injector ii distributes fuel from a point 9 radially outward from the centre of the whirl to take advantage of the whirl velocity in mixing the fuel and air. Figure 2 shows, diagrammatically, the manner in which even fuel distribution around the annulus may be effected; a plurality of fuel injectors 8a, 3, 8b are distributed evenly around the annulus of the injection zone, the fuel jets from each being directed circumferentially to diverge and intersect in the manner indicated by the arrows C; in this way a substantially even annulus of fuel is deposited in the region of the toroidal air whirl. This arrangement is considered to be most preferable for use in a combustion chamber of the annular type, anda practical example of such an application is illustrated in Figure 3 which represents a cross-section in a longitudinal plane through the axis of a chamber intended for use in a gas turbine power unit. With reference to Figure 3, an annular air casing I receives the output of air (indicated by the arrow A) from a compressor (not shovm) through an annulardischarge passage 2, and encloses an annular flame chamber 3 which is provided at its upstream end with three-'circumferentially 'extending sets of'air inlets, i, 5 andfi respectively,

disposed in a manner similar to that described with reference to Figure 1, to produce, inside the flame chamber, a toroidal whirl whose annulus is concentric with that of the flame chamber as indicated by the arrows B. Fuel injectors I are arranged to project into the chamber 3 and to discharge fuel in the circumferential direction in the zone of the whirl, but offset with respect to the annular axis about which the whirling takes place as described withreference to Figures 1 and 2. The air entering the inlets 4, 5 and 6 in the flame-chamber 3 constitutes the primary air for combustion and mixing, and secondary air is supplied through the ports 8,. 9, and II] respectively as indicated by the arrows C in the chamber downstream from the injection I zone. The products of combustion are finally ejected from the chamber through the annularoutlet II to the blade annulus of a turbine (not the direction of the arrow A through a part 2 and enclosing a flame tube 3. The upstream end of the flame tube 3 has a tubular centre piece 4 providing an axial passage extending into the tube, which centre piece has radial ports 5 therein to provide, in conjunction with ports 6 and I in the upstream end wall 8 and the radially outer wall 9 of the flame t'ube respectively, air jets producing a whirl as indicated by the arrows B. The upstream end I of the centre piece projects from the end wall 8 into the airstream to form a splitter allowing a metered quantity of air to enter the axial passage and diverting the remainder radially outwardly over the upstream end of the flame tube. The latter is provided with shroud II axially spaced from it which screens the axially directed air inlets 6 in the end wall 8 and meters off a proportion of the air flow diverted by the splitter I0 for supply through these inlets. The tubular centre piece 4 is extended axially beyond the radial air ports by which it contributes to the formation of the toroidal whirl in the form of a cone I2, tapering in the downstream direction and having holes I3 in it to supply secondary combustion air to the core of the flame. The balance of air supply as between this extension and the preceding radial air ports is determined by a tubular splitter I4 at the base of the cone I2. Additional secondary combustion air is supplied through holes I5 in the outer wall!) of the flame tube. Fuel is injected circumferentially into the region of the whirl from a plurality of nozzles I6 in a manner similar to that of the embodiments previously described.

Although the foregoing description relates to a self-contained combustion apparatus having a single flame tube it is apparent that an apparatus could comprise a plurality of flame tubes similar to that described arranged for operation in parallel in annular or other formation in a suitable air casing.

The invention may be applied with particular advantage in association with the use of gaseous or vaporised fuel, since the use of such fuel in itself assists in minimising the flame length, and the features of the invention accentuate this advantage, and also because multiple injection giving good fuel distribution is facilitated by the use of gas or vapour. I

I claim:

1. In a combustion apparatus for burning fuel the combination of a combustion chamber having an outlet for combustion products at one end and a chamber annulus at the other end comprising radially spaced inner and outer tubular walls defining a zone of substantially annular cross-section and an end wall substantially enclosing said zone at the end of the chamber remote from said outlet, means for introducing combustion air in said chamber annulus comprising a circumferential group of spaced air inlet openings in said inner wall located in a first common plane transverse to the axis of the chamber and a second circumferential group of spaced air inlet openings'in said outer wall located in a second common plane transverse to .the axis of the chamber and axially spaced from said first plane, said outer wall being characterized by the absence of openings in said first plane and said inner wall being characterized by the absence of openings in said second plane, means for supplying combustion air to said inlet openings in such a manner that discrete jets of air produced by the inlet openings of each respective group flow transversely across said chamber annulus and thence axially toward the jets produced by the inlet openings of the other group and being at least partly entrained therewith to describe a uniform symmetrical substantially toroidal path located substantially between said transverse planes containing the respective groups of inlet openings, and means for discharging fuel in said chamber annulus in a plane axially intermediate said first and second planes and substantially evenly around the circumference thereof of said annulus in a zone bounded by said toroidal path.

2. The combination according to claim 1, wherein said means for introducing combustion air in said chamber annulus further comprises a third circumferential group of spaced air inlet openings in said end wall located adjacent that one of said tubular walls whose circumferential group of inlet openings are contained in the more axially remote from said end wall of said two transverse planes, said means supplying combustion air to said first and second mentioned groups of inlet openings supplying air also to said third mentioned group of inlet openings in such a manner that discrete jets of air produced thereby flowaxially along said adjacent tubular wall.

References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 1,533,533 Wirrer Apr. 14, 1925 2,332,866 Muller Oct. 26, 1943 2,398,654 Lubbock et al Apr. 16, 1946 2,475,911 Nathan July 12, 1949 2,488,911 Hepburn et al. Nov. 22, 1949 2,510,571 Goddard H June 6, 1950 2,517,015 Mock et a1. Aug. 1, 1950 2,601,000 Nerad June 17, 1952

US2687010A 1947-11-03 1948-10-11 Combustion apparatus Expired - Lifetime US2687010A (en)

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Application Number Priority Date Filing Date Title
GB2926847A GB650462A (en) 1947-11-03 1947-11-03 Improvements in or relating to combustion apparatus

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US2687010A true US2687010A (en) 1954-08-24

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BE (1) BE485523A (en)
FR (1) FR973862A (en)
GB (1) GB650462A (en)
NL (1) NL72524C (en)

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2913874A (en) * 1955-03-30 1959-11-24 Gen Electric Tailpipe thrust augmentor
US2952126A (en) * 1955-05-10 1960-09-13 Midland Ross Corp Combustion unit for supplying hot gas for jet aircraft
DE1108516B (en) * 1956-04-03 1961-06-08 Bristol Siddeley Engines Ltd burning device
US3000183A (en) * 1957-01-30 1961-09-19 Gen Motors Corp Spiral annular combustion chamber
US3075352A (en) * 1958-11-28 1963-01-29 Gen Motors Corp Combustion chamber fluid inlet construction
US3080715A (en) * 1959-04-28 1963-03-12 Rolls Royce Combustion chamber
US3082603A (en) * 1955-10-28 1963-03-26 Snecma Combustion chamber with primary and secondary air flows
US3132484A (en) * 1960-05-18 1964-05-12 Rolls Royce Combustion products generator with diverse combustion and diluent air paths
US3355891A (en) * 1966-05-02 1967-12-05 Barry V Rhodes Ram jet engine and fuel injection system therefor
US3645095A (en) * 1970-11-25 1972-02-29 Avco Corp Annualr combustor
US4891936A (en) * 1987-12-28 1990-01-09 Sundstrand Corporation Turbine combustor with tangential fuel injection and bender jets
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5195315A (en) * 1991-01-14 1993-03-23 United Technologies Corporation Double dome combustor with counter rotating toroidal vortices and dual radial fuel injection
WO1995032395A1 (en) * 1994-05-25 1995-11-30 Westinghouse Electric Corporation Gas turbine combustor
WO2009056425A2 (en) * 2007-11-02 2009-05-07 Siemens Aktiengesellschaft A combustor for a gas-turbine engine

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
NL79375C (en) * 1951-05-31
US4955201A (en) * 1987-12-14 1990-09-11 Sundstrand Corporation Fuel injectors for turbine engines
US5165226A (en) * 1991-08-09 1992-11-24 Pratt & Whitney Canada, Inc. Single vortex combustor arrangement

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1533533A (en) * 1922-06-06 1925-04-14 Int Motor Co Combustion chamber for oil-burning furnaces
US2332866A (en) * 1937-11-18 1943-10-26 Muller Max Adolf Combustion chamber for gas-flow engines
US2398654A (en) * 1940-01-24 1946-04-16 Anglo Saxon Petroleum Co Combustion burner
US2475911A (en) * 1944-03-16 1949-07-12 Power Jets Res & Dev Ltd Combustion apparatus
US2488911A (en) * 1946-11-09 1949-11-22 Surface Combustion Corp Combustion apparatus for use with turbines
US2510571A (en) * 1946-05-11 1950-06-06 Esther C Goddard Combustion chamber with annular target area
US2517015A (en) * 1945-05-16 1950-08-01 Bendix Aviat Corp Combustion chamber with shielded fuel nozzle
US2601000A (en) * 1947-05-23 1952-06-17 Gen Electric Combustor for thermal power plants having toroidal flow path in primary mixing zone

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1533533A (en) * 1922-06-06 1925-04-14 Int Motor Co Combustion chamber for oil-burning furnaces
US2332866A (en) * 1937-11-18 1943-10-26 Muller Max Adolf Combustion chamber for gas-flow engines
US2398654A (en) * 1940-01-24 1946-04-16 Anglo Saxon Petroleum Co Combustion burner
US2475911A (en) * 1944-03-16 1949-07-12 Power Jets Res & Dev Ltd Combustion apparatus
US2517015A (en) * 1945-05-16 1950-08-01 Bendix Aviat Corp Combustion chamber with shielded fuel nozzle
US2510571A (en) * 1946-05-11 1950-06-06 Esther C Goddard Combustion chamber with annular target area
US2488911A (en) * 1946-11-09 1949-11-22 Surface Combustion Corp Combustion apparatus for use with turbines
US2601000A (en) * 1947-05-23 1952-06-17 Gen Electric Combustor for thermal power plants having toroidal flow path in primary mixing zone

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2913874A (en) * 1955-03-30 1959-11-24 Gen Electric Tailpipe thrust augmentor
US2952126A (en) * 1955-05-10 1960-09-13 Midland Ross Corp Combustion unit for supplying hot gas for jet aircraft
US3082603A (en) * 1955-10-28 1963-03-26 Snecma Combustion chamber with primary and secondary air flows
DE1108516B (en) * 1956-04-03 1961-06-08 Bristol Siddeley Engines Ltd burning device
US3088281A (en) * 1956-04-03 1963-05-07 Bristol Siddeley Engines Ltd Combustion chambers for use with swirling combustion supporting medium
US3000183A (en) * 1957-01-30 1961-09-19 Gen Motors Corp Spiral annular combustion chamber
US3075352A (en) * 1958-11-28 1963-01-29 Gen Motors Corp Combustion chamber fluid inlet construction
US3080715A (en) * 1959-04-28 1963-03-12 Rolls Royce Combustion chamber
US3132484A (en) * 1960-05-18 1964-05-12 Rolls Royce Combustion products generator with diverse combustion and diluent air paths
US3355891A (en) * 1966-05-02 1967-12-05 Barry V Rhodes Ram jet engine and fuel injection system therefor
US3645095A (en) * 1970-11-25 1972-02-29 Avco Corp Annualr combustor
EP0349635A4 (en) * 1987-12-28 1990-05-14 Sundstrand Corp Turbine combustor with tangential fuel injection and bender jets.
EP0349635A1 (en) * 1987-12-28 1990-01-10 Sundstrand Corp Turbine combustor with tangential fuel injection and bender jets.
US4891936A (en) * 1987-12-28 1990-01-09 Sundstrand Corporation Turbine combustor with tangential fuel injection and bender jets
US5109671A (en) * 1989-12-05 1992-05-05 Allied-Signal Inc. Combustion apparatus and method for a turbine engine
US5195315A (en) * 1991-01-14 1993-03-23 United Technologies Corporation Double dome combustor with counter rotating toroidal vortices and dual radial fuel injection
WO1995032395A1 (en) * 1994-05-25 1995-11-30 Westinghouse Electric Corporation Gas turbine combustor
US5636510A (en) * 1994-05-25 1997-06-10 Westinghouse Electric Corporation Gas turbine topping combustor
WO2009056425A2 (en) * 2007-11-02 2009-05-07 Siemens Aktiengesellschaft A combustor for a gas-turbine engine
WO2009056425A3 (en) * 2007-11-02 2010-06-24 Siemens Aktiengesellschaft A combustor for a gas-turbine engine
US20100293953A1 (en) * 2007-11-02 2010-11-25 Siemens Aktiengesellschaft Combustor for a gas-turbine engine
RU2478879C2 (en) * 2007-11-02 2013-04-10 Сименс Акциенгезелльшафт Combustion assembly for gas turbine engine
US8984889B2 (en) 2007-11-02 2015-03-24 Siemens Aktiengesellschaft Combustor for a gas-turbine engine with angled pilot fuel nozzle

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BE485523A (en) grant
NL72524C (en) grant
GB650462A (en) 1951-02-28 application
FR973862A (en) 1951-02-15 grant

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