US3126705A - Combustion system - Google Patents

Combustion system Download PDF

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US3126705A
US3126705A US3126705DA US3126705A US 3126705 A US3126705 A US 3126705A US 3126705D A US3126705D A US 3126705DA US 3126705 A US3126705 A US 3126705A
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disc
compressor
fuel
combustion chamber
turbine
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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
    • F23R3/38Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising rotary fuel injection means
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00Stators
    • F01D9/02Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04Air intakes for gas-turbine plants or jet-propulsion plants
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/16Cooling of plants characterised by cooling medium
    • F02C7/18Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/22Fuel supply systems
    • F02C7/236Fuel delivery systems comprising two or more pumps
    • F02C7/2365Fuel delivery systems comprising two or more pumps comprising an air supply system for the atomisation of fuel
    • 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
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies, e.g. for aircraft

Description

March '31, 1964 R p PRQBERT ETAL 3,126,705
COMBUSTION SYSTEM FIG. 5.
pw WM 1%:
3 Sheets-Sheet 2 COMBUSTION SYSTEM R. P. PROBERT ETAL March 31, 1964 Filed` March 22, 1957 March 31, 1964 R, p, PROBERT ETAL 3,126,705
COMBUSTION SYSTEM Filed March 22. 1957 3 Sheets-Sheet 5 Flg 6 Fig. 4.
United States Patent O 3,126,705 COMBUSTION SYSTEM Rhys Price Probert, Farnborough, Peter Martin, Farnham, and Edward Langford Hartley, Farnborough, England, assignors, by mesne assignments, .to Power Jets (Research and Development) Limited, London, England, a British company Filed Mar. 22, 1957, Ser. No. 647,918 Claims priority, application Great Britain Mar. 26, 1956 3 Claims. (Cl. 60-39.74)
This invention relates to a combustion method and to combustion apparatus particularly, but not exclusively, for u'se in gas turbine plant. ln Conventional gas turbine plant in which a compressor discharges a stream of air into a combustion Chamber wherein fuel is burnt in the air stream and the combustion gases so produced are discharged through a turbine which drives the compressor, the length of the fiow path between the compressor and the turbine is determined partly by the means used for atomisation of the fuel and mixing of the fuel with the air stream in the combustion Chamber, and partly by the degree of ditfusion required between the compressor air outlet and the combustion Chamber air inlet. If such diffusion could be dispensed with a reduction in the length of the said fiow path could be made. This would however involve an air Velocity at the combustion Chamber inlet which is impracticable with the combustion apparatus in general use at present.
The present invention provides a combustion method which comprises discharging a Sheet of liquid fuel trans- Versely into a high Velocity air stream so that the sheet of fuel is shattered thereby into droplets which are entraned in the air stream, and subsequently burning the resulting mixture of fuel and air. According to further features of the invention, the Sheet of fuel is discharged radially outwards into a high Velocity annular air stream, and the mixture of fuel and air is burnt in a gas path extending in a direction substantially normal to the plane of discharge of the fuel.
The present invention also provides a combustion chamber having a fuel-carrying member rotatably mounted in the upstream end thereof, means for rotating said member and means for introducing liquid fuel on to the surface of said member so theat the fuel is thrown off said member as a sheet of fuel.
Again the present invention provides a gas turbine plant comprising a dynamic compressor, a turbine, an annular combustion Chamber disposed between the compressor and the turbine, a shaft extending along the aXis of the combustion Chamber and Coupling the turbine to drive the compressor, a fuel-Carrying disc mounted for rotation with said shaft, the periphery of the disc being spaced from a boundary wall of the combustion space within the Chamber to provide an annular inlet for air from the compressor to said space, and means for introducing liquid on to a side face of the disc so that the fuel is thrown off the disc into the airstream entering the combustion space.
In order that the fuel shall adhere to and spread evenly over the surface of the rotating disc under Centrifugal force before discharging from the disc, it has been found necessary to use a disc of which at least a radially outer portion of the fuel-Carrying surface is slightly inwardly Coned. The means for admitting fuel on to the surface of the disc are desiflned to admit such a quantity of fuel with respect to the speed of the disc that a sheet of fuel, as distinct from a spray of fuel droplets, is discharged from the disc surface. The quantity of fuel which can be discharged as a sheet in this way is several hundred times greater than that which could be discharged under the same conditions if a spray of fuel droplets were required.
ICC
In order to provide a high Velocity air stream at the combustion Chamber inlet, the compressor is arranged to discharge air substantially directly, without diffusion into the combustion Chamber inlet. The omission of the normal diffuser between the compressor outlet and the combustion Chamber inlet thus shortens the length of the flow path. As the high Velocity air stream sweeps over the periphery of the fuel-Carrying disc, the Sheet of fuel is shattered into small droplets which are entrained in the air stream to provide a gaseous combustible miXture which is then ignited.
Fuel may be admitted to either face of the fuel-carrying disc, but if the downstream facing surface of the disc is exposed to the heat of the combustion products within the Chamber, the fuel will normally be admitted on to this surface in order to Cool the surface.
The Sudden expansion of the compressor air into the combustion Chamber produces violent turbulence in the combustible mixture and this greatly asssts flame stabilisation in the combustion zone immediately downstream of the disc. In order to increase this turbulence and thereby assist combustion of the gaseous combustible mixture, paddles are secured to the rearward surface of the disc.
in most Conventional forms of gas turbine plant the combustion Chamber comprises a perforated fiame tube and an outer wall which forms with the flame tube a duct for secondary air. The secondary air flowing in this duct passes through the perforations in the flame tube along its length and mixes with the combustion gases within the flame tube. The length of such combustion chamber is therefore dependent at least partly on the degree of mixing of gases and air required before they enter the turbine. According to a further feature of the present invention, the combustion Chamber has a substantially imperforate flame tube surrounded by a casing and secondary air fiowing between the flame tube and the Casing is mixed with the combustion gases between the inlet guide vanes of the turbine; in this way a reduction may be effected both in the length of the combustion chamber and in the pressure losses within the Chamber. In addition, by directing the secondary air on to these inlet guide vanes, Cooling of the vanes is effected.
Two specific embodiments of the present invention as applied to a gas turbine jet Propulsion plant are shown in the accompanying drawings in which:
FIGURE l is a diagrammatic side view of a gas turbine jet propulsion plant.
FIGURE 2 is a longitudinal half-sectional View showing in detail part of the gas turbine plant of FIGURE 1.
FIGURES 3, 4 and 5 are Sections on the lines III-III, IV-IV and V-V respectively of FIGURE 2 and FIGURE 6 is a view, Corresponding to the View shown in FIGURE 2, of an alternative form of combustion means.
FIGURE 7 is a section on the line VII- VII of FIG- URE 6.
The gas turbine shown in FIGURE 1 has an axialflow compressor 1 receiving air from atmosphere through an annular intake 2, an annular combustion Chamber 3 into which the compressor discharges air under compression, an aXial-flow turbine 4 connected to receive Combustion gases from the combustion Chamber and secondary air from the compressor, and a jet pipe 5 through which the turbine exhaust gases discharge to atmosphere as a propulsive jet. The compressor comprises stator blades 6 supported on a stator casing 7 and rotor blades 3 supported on a rotor body 9, and similarly the turbine comprises stator blades l supported on a stator casing 11 and rotor blades 12 supported on a rotor disc 13, the compressor rotor body and the turbine rotor disc being connected by a shaft 14 which extends along the aXis of the plant. This rotor structure is supported on the axis of the plant on a front bearing 15, a rear bearing 16 and in certain instances by an intermediate bearing 17, although it is envisaged that by suitable dimensioning of components this intermediate bearing may be omitted. Fuel is introduced into the combustion chamber from a fuel tank 18 through a control valve 19, a pump 20, a first fuel pipe 21 which extends across the air intalre and terminates in a rotary seal 22, a second fuel pipe 23 which extends from the rotary seal axially through the compressor rotor, and a fuel-carrying disc 24 mounted for rotation with the shaft on to the surface of which fuel is conducted from the second pipe.
As is shown in greater detail in FIGURE 2, the combustion chamber is formed by an outer casing 30 and a fiame tube 31 having outer and inner fiame tube walls 32, 33, the inlet end of the combustion chamber, which is closely adjacent to the outlet of the compressor, forming an annular substantially non-diffusing, i.e. Constant velocity, passage defined between the outer casing 30 and an inner wall 34 supported by the radially inner end of a row of compressor outlet stator blades 6. The fiame tube outer wall 32 divides the airflow from the compressor into two annular streams, one fiowing over the outside of the fiame tube through a duct 35 formed between the wall 32 and the casing 30 and the other fiowng into the fiame tube. The air inlet to the fiame tube is divided into a large number of individual streams by means of fingers 36 which extend transversely across the inlet and serve to produce an even flow of air into the fiame tube. The outer wall of the fiame tube is substantially imperforate, that is to say the wall is either imperforate over its entire length or provided with only a small number of holes or slits to admit a small quantity of air from the duct 35 sufficient for example to form a protective layer of cooling air over the inner surface of this wall.
At the rearward end of the compressor, the compressor rotor is formed with a rearwardly extending stub shaft 37 which is journalled in the intermediate bearing 17 and rearward of the bearing is provided with splines 33 which engage the forward end of the shaft 14. A truncated conical support member 39 mounted on the stub shaft between the splines and the bearing is provided with a radial flange 40 at its rearward end, and the fuel disc 24 is held in position on this fiange by a second similar support member 41 flanged at 42 which also is mounted on the shaft, the disc being secured between the two fianges by bolts 43, and so dimensioned that its periphery, which is spaced from the outer wall of the flame tube, is swept by the annular air stream entering the flame tube. A fuel distributing plug 44 provided with radial passages 45 is threaded into the hollow interior of the stub shaft 37 and is connected to the rearward end of the second fuel pipe 23 so that by way of apertures 46 in the wall of the stub shaft, a path for fuel is formed from the pipe 23 to a space 47 between the support members 39, 41. Radial slots 48 formed in the forward face of the flange 42 allow fuel from the space 47 to flow into an annular trough 49 formed in the rearward face of the disc 24, from which under centrifugal force it spills on to the exposed radially outer portion of the rearward surface of the disc. This exposed surface portion is slightly inwardly coned, the angle which the coned portion of the surface makes with a radial plane being not less than 11/2, preferably between 3 and so that the fuel is caused to adhere to the disc surface by centrifugal force as it spreads radially outwards over the disc. Since the rearward or downstream facing surface of the disc forms the upstream boundary wall of the combustion space within the fiame tube, the fuel flowing radially outwards over this surface serves to protect the surface from the heat of the combustion gases within the fiame tube.
The apertures 46 in the stub shaft and the dimensions of the fuel exit through the trough 49 are so arranged ,4 that over at least part of the range of Operating speeds of the shaft and consequently of the fuel disc, such quantity of fuel may be carried by the disc as will produce a sheet of fuel as distinct from a fuel spray thrown off centrifugally from the disc into the incoming high Velocity air stream so that a high degree of atomisation is effected. One or more igniters 50 extend through the outer wall of the fiame tube within the fuel and air mixng zone downstream of the fuel-carrying disc in which primary combustion takes place. In order to assist combustion of the fuel/ air mixture in this primary combustion zone by increasing the natural turbulence of the mixture, and thereby producing a shortening of the fiame length within the fiame tube, paddles 51 are mounted on the rearward face of the fuel disc for rotation therewith as shown more clearly in FIGURE 3. The angle which these paddles make with a radius of the disc is about 30, and the leading edge of each paddle with respect to the direction of rotation of the disc (indicated by the arrow in FIG. 3) is spaced radially inwardly with respect to the trailing edge.
The downstream end of the fiame tube outer wall 32 is corrugated to form a number of longitudinally extending chutes 56 which are equally spaced apart circumferentially and accordingly form circumferentially alternating passages 57, 58 for hot gases from the interior of the fiame tube and for secondary air from the duct formed between the outer wall of the fiame tube and the casing, respectively. The chutes extend across the full width of the combustion chamber outlet but the passages formed by the chutes may be sub-divided by thin vanes 59 into a number of individual passages to assist in producing a smooth gas flow. The sides of the chutes may be parallel, but preferably they converge towards the axis of the plant as shown in FIGURE 4.
The inlet guide vanes 10 of the turbine are located across the combustion chamber Outlet path immediately downstream of the outlet of the chutes and as shown in FIGURE 5, each chute is so aligned with respect to a turbine inlet guide vane as to cause the secondary air to flow mainly over the convex surface of the guide vane, only sufficient being directed on to the concave surface as is needed to distribute a thin film of air over the surface for cooling purposes. Some of the secondary air flowing over the convex surface of a vane gravitates off towards the concave surface of the next adjacent guide Vane and in so doing mixes with the combustion gases in the guide vane passages. By means of this arrangement, good mixng of the secondary air and combustion gases is eifected without destroying the general axial flow of each, and in consequeence the pressure losses in mixng are low.
The inner wall 33 of the fiame tube is of hollow construction to provide a passage 60 through the hollow interior of the wall for cooling air, and the radially inner ends of the secondary air passages 58 are placed in communication with the passage 60 by means of apertures 62 in the wall 33 at the downstream end of the fiame tube which form the inlet to the passage so that some of the secondary air is metered into this passage and floWs in the forward direction to an outlet 63 at the upstream end of the fiame tube. A small fan 61 in the form of a centrifugal compressor rotor is mounted at this outlet for rotation With the disc 24 to create a region of low pressure at the Outlet which serves to draw secondary air through the passage 60 and to discharge this air into the primary combustion zone within the fiame tube.
FIGURE 6 shows diagrammatically a modification of the embodiment of FIGURE 2 in which a fuel-carrying disc is supported directly on a shaft 71 connecting the turbine rotor 13 to drive the compressor rotor 9, and a separate disc 72 also supported on the shaft 71 is provided, on the downstream side of the fuel disc, on which turbulence-forming paddles 73 are carried. Fuel is admitted to the space between the two discs through apertures 74 in the shaft, and flows into an annular trough 75 formed in the surface of the disc 70 from which it is distributed over the disc surface.
FIGURE 6 also shows a modified form of flame tube inlet in which fingers 76 of triangular lShape extend radially inwards across this inlet terminating short of the wall 34; by this arrangement the air flow adjacent the wall 34 is uninterrupted and thus enters the flame tube as a truly annular stream. As shown in FIGURE 7, the passages between the fingers 76 are tapered and expand in the radially inward direction so that the air flow is rnetered into the flame tube, increasing in the radially inward direction. Either or both of these features may equally be applicable to the fingers 36 of the first embodi' ment.
We clairn:
1. A gas turbine comprising a compressor including a rotor; a turbine mounted coaxially with and axially spaced from the compressor and including a rotor; an annular combustion Chamber lying between the compressor and turbine and comprising coaxial inner and outer generally axially extending walls defining between them a flow path from the compressor to the turbine; a shaft drivingly connecting the compressor and turbine rotors and extending coaxially within said inner wall; a disc mounted on the shaft for rotation therewith at the compressor end of the combustion Chamber, said disc having a periphery spaced outwardly from said inner wall but spaced inwardly from said outer wall to define with the latter a generally aXially facing inlet to the combustion chamber, said disc also having a side surface extending generally transversely to the axial direction; means defining a substantially Constant Velocity passage connecting said compressor to said combustion chamber inlet; and means for introducing liquid fuel onto said side surface of the disc at a position spaced inwardly from the periphery at such a rate in relation to the speed of rotation of the disc that the fuel is thrown off the periphery of the disc across the inlet as a Sheet of fuel.
2. A gas turbine according to claim 1 wherein said side surface is the downstream face of the disc in relation to the direction of flow through the combustion Chamber and partly defines the compressor end of the combustion Chamber.
3. A gas turbine according to claim 1 wherein at least a radially outer portion of said side surfacee of the disc is inwardly coned.
References Cited in the file of this patent UNITED STATES PATENTS 762,048 Gibbs June 7, 1904 1,996,336 Junkers Apr. 2, 1935 2,086,377 Bryan July 6, 1937 2,249,878 Asbury July 22, 1941 2,284,141 Funk May 26, 1942 2,536,600 Goddard Jan. 2, 1951 2,547,959 Miller Apr. 10, 1951 2,568,921 Kroon Sept. 25, 1951 2,595,505 Bachle May 6, 1952 2,602,292 Buckland et al. July 8, 1952 2,657,531 Pierce Nov. 3, 1953 2,659,196 Brown Nov. 17, 1953 2,694,291 Rosengart Nov. 16, 1954 2,705,401 Allen et al. Apr. 5, 1955 2,720,750 Schelp Oct. 18, 1955 2,801,519 Wood Aug. 6, 1957 2,929,2()9 Schirmer Mar. 22, 1960 FOREIGN PATENTS 114,844 Switzerland July 1, 1926

Claims (1)

1. GAS TURBINE COMPRISING A COMPRESSOR INCLUDING A ROTOR; A TURBINE MOUNTED COAXIALLY WITH AND AXIALLY SPACED FROM THE COMPRESSOR AND INCLUDING A ROTOR; AN ANNULAR COMBUSTION CHAMBER LYING BETWEEN THE COMPRESSOR AND TURBINE AND COMPRISING COAXIAL INNER AND OUTER GENERALLY AXIALLY EXTENDING WALLS DEFINING BETWEEN THEM A FLOW PATH FROM THE COMPRESSOR TO THE TURBINE; A SHAFT DRIVINGLY CONNECTING THE COMPRESSOR AND TURBINE ROTORS AND EXTENDING COAXIALLY WITHIN SAID INNER WALL; A DISC MOUNTED ON THE SHAFT FOR ROTATION THEREWITH AT THE COMPRESSOR END OF THE COMBUSTION CHAMBER, SAID DISC HAVING A PERIPHERY SPACED OUTWARDLY FROM SAID INNER WALL BUT SPACED INWARDLY FROM SAID OUTER WALL TO DEFINE WITH THE LATTER A GENERALLY AXIALLY FACING INLET TO THE COMBUSTION CHAMBER, SAID DISC ALSO HAVING A SIDE SURFACE EXTENDING GENERALLY TRANSVERSELY TO THE AXIAL DIRECTION; MEANS DEFINING A SUBSTANTIALLY CONSTANT VELOCITY PASSAGE CONNECTING SAID COMPRESSOR TO SAID COMBUSTION CHAMBER INLET; AND MEANS FOR INTRODUCING LIQUID FUEL ONTO SAID SIDE SURFACE OF THE DISC AT A POSITION SPACED INWARDLY FROM THE PERIPHERY AT SUCH A RATE IN RELATION TO THE SPEED OF ROTATION OF THE DISC THAT THE FUEL IS THROWN OFF THE PERIPHERY OF THE DISC ACROSS THE INLET AS A SHEET OF FUEL.
US3126705D 1956-03-26 Combustion system Expired - Lifetime US3126705A (en)

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Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB980363A (en) * 1961-12-04 1965-01-13 Jan Jerie Improvements in or relating to gas turbines
GB2030653B (en) * 1978-10-02 1983-05-05 Gen Electric Gas turbine engine combustion gas temperature variation
US4733538A (en) * 1978-10-02 1988-03-29 General Electric Company Combustion selective temperature dilution
US4739621A (en) * 1984-10-11 1988-04-26 United Technologies Corporation Cooling scheme for combustor vane interface
US4769996A (en) * 1987-01-27 1988-09-13 Teledyne Industries, Inc. Fuel transfer system for multiple concentric shaft gas turbine engines

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US762048A (en) * 1903-07-16 1904-06-07 Henry Berg Fuel-burner.
CH114844A (en) * 1924-10-23 1926-07-01 Sulzer Ag Oil burner with rotating atomizer.
US1996336A (en) * 1931-01-08 1935-04-02 Firm Junkers & Co G M B H Burner for liquid fuel
US2086377A (en) * 1931-09-23 1937-07-06 Bryan Steam Corp Oil burner
US2249878A (en) * 1939-01-21 1941-07-22 Charles T Asbury Oil burner
US2284141A (en) * 1940-07-25 1942-05-26 Advance Aluminum Castings Corp Suction fan unit
US2536600A (en) * 1948-02-07 1951-01-02 Daniel And Florence Guggenheim Rotating, feeding, and cooling means for combustion chambers
US2547959A (en) * 1948-01-27 1951-04-10 Westinghouse Electric Corp Centrifugal fuel feeding system for annular combustion chambers
US2568921A (en) * 1948-04-27 1951-09-25 Westinghouse Electric Corp Combustion chamber with rotating fuel nozzles
US2595505A (en) * 1946-04-20 1952-05-06 Continental Aviat & Engineerin Coaxial combustion products generator, turbine, and compressor
US2602292A (en) * 1951-03-31 1952-07-08 Gen Electric Fuel-air mixing device
US2657531A (en) * 1948-01-22 1953-11-03 Gen Electric Wall cooling arrangement for combustion devices
US2659196A (en) * 1949-08-09 1953-11-17 United Aircraft Corp Centrifugal fuel supply means for jet engine afterburners
US2694291A (en) * 1948-02-07 1954-11-16 Henning C Rosengart Rotor and combustion chamber arrangement for gas turbines
US2705401A (en) * 1950-12-02 1955-04-05 Armstrong Siddeley Motors Ltd Vaporising means for liquid fuel combustion chambers
US2720750A (en) * 1947-11-04 1955-10-18 Helmut R Schelp Revolving fuel injection system for jet engines and gas turbines
US2801519A (en) * 1951-02-17 1957-08-06 Garrett Corp Gas turbine motor scroll structure
US2929209A (en) * 1953-09-15 1960-03-22 Phillips Petroleum Co Combustion of fuel on the surface of a rotating disc

Patent Citations (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US762048A (en) * 1903-07-16 1904-06-07 Henry Berg Fuel-burner.
CH114844A (en) * 1924-10-23 1926-07-01 Sulzer Ag Oil burner with rotating atomizer.
US1996336A (en) * 1931-01-08 1935-04-02 Firm Junkers & Co G M B H Burner for liquid fuel
US2086377A (en) * 1931-09-23 1937-07-06 Bryan Steam Corp Oil burner
US2249878A (en) * 1939-01-21 1941-07-22 Charles T Asbury Oil burner
US2284141A (en) * 1940-07-25 1942-05-26 Advance Aluminum Castings Corp Suction fan unit
US2595505A (en) * 1946-04-20 1952-05-06 Continental Aviat & Engineerin Coaxial combustion products generator, turbine, and compressor
US2720750A (en) * 1947-11-04 1955-10-18 Helmut R Schelp Revolving fuel injection system for jet engines and gas turbines
US2657531A (en) * 1948-01-22 1953-11-03 Gen Electric Wall cooling arrangement for combustion devices
US2547959A (en) * 1948-01-27 1951-04-10 Westinghouse Electric Corp Centrifugal fuel feeding system for annular combustion chambers
US2694291A (en) * 1948-02-07 1954-11-16 Henning C Rosengart Rotor and combustion chamber arrangement for gas turbines
US2536600A (en) * 1948-02-07 1951-01-02 Daniel And Florence Guggenheim Rotating, feeding, and cooling means for combustion chambers
US2568921A (en) * 1948-04-27 1951-09-25 Westinghouse Electric Corp Combustion chamber with rotating fuel nozzles
US2659196A (en) * 1949-08-09 1953-11-17 United Aircraft Corp Centrifugal fuel supply means for jet engine afterburners
US2705401A (en) * 1950-12-02 1955-04-05 Armstrong Siddeley Motors Ltd Vaporising means for liquid fuel combustion chambers
US2801519A (en) * 1951-02-17 1957-08-06 Garrett Corp Gas turbine motor scroll structure
US2602292A (en) * 1951-03-31 1952-07-08 Gen Electric Fuel-air mixing device
US2929209A (en) * 1953-09-15 1960-03-22 Phillips Petroleum Co Combustion of fuel on the surface of a rotating disc

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