US2575264A - Construction and mounting of fuel burner nozzles, especially for gas turbines and like apparatus - Google Patents

Construction and mounting of fuel burner nozzles, especially for gas turbines and like apparatus Download PDF

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Publication number
US2575264A
US2575264A US641145A US64114546A US2575264A US 2575264 A US2575264 A US 2575264A US 641145 A US641145 A US 641145A US 64114546 A US64114546 A US 64114546A US 2575264 A US2575264 A US 2575264A
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United States
Prior art keywords
burner
fuel
mounting
duct
construction
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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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US641145A
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English (en)
Inventor
Feilden Geoffrey Bertra Robert
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Power Jets Research and Development Ltd
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Power Jets Research and Development Ltd
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/062—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with aft fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • Y—GENERAL 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
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention relates to combustion apparatus more especially of a combustion gas turbine, in which fuel is burnt in a flame tube enclosed within an air duct carrying a flow of combustion-supporting air past the flame tube from a point upstream to a point downstream thereof, the fuel being injected into the flame tube by a burner mounted for downstream injection relative to said flow.
  • the invention is concerned particularly with the construction and arrangement of fuel burner nozzles, sometimes called injectors, in such apparatus.
  • the engine is one having a multistage axial flow compressor with an annular diffuser outlet leading into an annular combustion chamber containing a plurality of axially directed flame tubes or other combustion elements, the outlet of the combustion chamber being united direct to the nozzle ring of the turbine, the whole comprising a gas turbine aero engine which in general form is symmetrical about its own main axis.
  • the upstream end of the flame tube has a focal point of attachment at or in the region of the burner nozzle.
  • a class of burner to which the invention particularly relates is that in which a vortex chamber is employed in which the fuel is caused to swirl and to be ejected through a small orifice from which it sprays with fine atomization.
  • the invention is concerned with the construction of burner, its removability, its mounting and its associated ducting and it is an object of the invention to provide a construction in which a burner, being itself reasinably simple, compact and easily made, can be removed and replaced with a minimum of preliminary movement away from its support and which when mounted is supported robustly so that it can if necessary be used for the further support of a flame tube or its equivalent, whilst the necessity is avoided for specially provided unions with its fuel supply ducting extraneous to its mounting.
  • a burner according to the invention is therefore capable of being mounted direct on part of the engine structure in an internal and normally inaccessible location.
  • the internal mounting of a burner in a gas turbine engine presents many difllculties which the invention seeks to overcome, but at the same time such *intemal mounting has in some designs of engine great inherent advantage.
  • the burner body is detachably mounted for downstream injection on a structural member which extends across the ducting carrying the gas stream and divides such stream, the fuel supply connection to the burner passing through the said structural member and the burner body forming a point of support for the upstream end of the flame tube into which the fuel is injected, said support being such as to allow axial withdrawal of said flame tube from the burner mounting, and including releasable locking means adapted to retain said flame tube against involuntary displacement. It is intended that the burner should be releasably coupled to the fuel supply connection by means which allow the removal of such body independently of its fuel supply connection.
  • a preferred embodiment of the invention is one in which appropriate fuel ducting opens at a face or platform on the structural member, and the burner body is applied to said face, the fuel ducting in the burner body registering with that of the structural member, and the construction being such that sealed fluid-flow engagement of the respective ducts is effected by the act of mounting the body on the structural member, without involving pipe unions.
  • the structural element upon which the burner is mounted is a partition dividing radially an annular duct, for example, the annular outlet duct of an axial flow compressor, there being a number of such partitions spaced around the axis of the compressor to deflne axial diffuser passages and each carrying a burner.
  • Figure 1 is a part sectional elevation of the forward part of a jet propulsion gas turbine power 3 unit of a design to which the invention has particular application;
  • Figure 2 is a fragmentary axial section through the outlet annulus of a compressor such as that illustrated in Figure 1, showing the mounting of a fuel burner and a part of an associated combustion casing or flame tube;
  • Figures 3 and 4 are sections respectively on lines 3-3 and 4-4 of Figure 2;
  • Figure 5 is a view similar to Figure 2, but illustrates a different type of burner and mounting.
  • the principal elements are an axial flow compressor generally indicated as I supplying air by way of an outlet annulus generally indicated at 2 to an annular combustion arrangement generally indicated as 3, from which the products of combustion and air pass to an axial flow turbine generally indicated as 6.
  • the compressor which is conventionally illustrated since its precise details of construction are not relevant to an understanding of the present invention, comprises a bladed rotor la and a bladed stator lb, of which the former is connected by a shaft I c with a rotor 6a of the turbine *6 to be driven thereby.
  • the axial flow turbine 6 has rotary nozzle blading 6b carried by a rotor 60 disposed upstream of the rotor 6a and supported through bearing means Ed on a stationary bearing sleeve lie within which the shaft is has its rear bearing.
  • the nozzleblading 62) carries, radially outwardly, a rotary shroud 61 supporting the interstage guide blading of the turbine.
  • the turbine exhausts by a duct 69 terminating in a propulsive J'et (not shown) which is surrounded concentrically by an augmentor duct 6h in which runs thrust augmenting blading 6i carried by the nozzle blading 6b.
  • the combustion arrangement 3 comprises inner and outer walls 3a, 3b together forming an annular air casing connecting the compressor outlet 2 and the turbine i.
  • annular air casing In the annular air casing are arranged individual flame tubes 4 in which combustion actually takes place, a part of the total airflow being admitted for that purpose through suitable ports in the flame tubes, and the remainder of the air passing around the outside of the flame tubes to mix with the products of combustion at the outlets therefrom.
  • Each flame tube 5 is supported at its downstream end so as to be capable of some rearward axial movement by telescoping with a fixed outlet chute la supported as a part of the main structure by means of a hollow strut 4b, and at its upstream end on a fuel injection nozzle or burner 5, the last mentioned support being shown conventionally in Figure 1 and in greater detail in alternative constructional forms in Figures 2-5.
  • the fuel burners 5 are mounted each at the downstream edge of a radial partition or splitter 1 extending axially in the compressor outlet annulus 2 (that is to say, in the same general direction as the air stream), there being several of such partitions defining axial outlet channels and each mounting a burner.
  • Each splitter is of substantially aerofoil form in axial cross section and has a radial passage 8 which not only accommodates a pipe connection In for the external supply, but also constitutes a. passage for the flow of cooling air be tween the outer atmosphere and the inner chamber enclosed by the compressor outlet annulus 2-
  • One or more further passages ll are drilled between the radial passage 8 and the downstream pipe connection i0.
  • the ducting system of the burner is of the kind in which there are tangential swirl ports 18 connected to a swirl chamber I9 having a discharge nozzle 20 and spill duct 2! for the return of surplus fuel to the supply.
  • the nozzle 20 is formed in a separate end plate held in position by a sleeve nut 22 having air passages 23 therethrough and provided with a part-spherical outer rim mating with a corresponding seating of a screwed sleeve 24 to which the upstream end of a flame tube 4 is attached, the spherical seating of said sleeve 24 being held in engagement with that of the burner by a threaded'locking sleeve 25 which is dimensioned and shaped to form an air passage at 26 between itself and the burner body and com-'- municating with the passages 23, so as to supply air within the flame tube 4 in the vicinity of the discharge nozzle 20.
  • the sleeve 25 is secured by a wire fastener 25a.
  • the trailing edge of the splitter is bossed at 21 to provide a plane face for the attachment of the burner body.
  • the main body of the burner is cylindrical, its supporting base is of flattened cross section and the fuel ducts and bolt holes formed in a line along it corresponding to the trailing edge of the splitter, so that a minimum of aerodynamic interference 'may be achieved.
  • the burner support comprises a ferrule 30 which is locked in an axial duct ll in the splitter by threaded engagement at one end with the pipe connection In and at the other with a lock nut 3
  • the burner body and the ferrule have complementary part spherical seatings which are held in engagement by a union nut 32, whilst the flame tube mounting sleeve 24 is in this case threaded externally for engagement by a securing nut 34 which holds the sleeve 24 against a seating flange 33 on the burner body.
  • the sleeve 24 also provides the air passage 28 to the region of the burner nozzle.
  • a combustion apparatus comprising a duct carrying a combustion-supporting air flow, a structural member attached at its ends to opposed walls of the duct and extending across the duct to divide the air flow, through which member passes a fuel supply passage, a flame tube enclosed by and extending longitudinally of said duct and located downstream of said member, a fuel injection burner receiving fuel from said supply passage detachably mounted on the downstream side of said member and having a body extending downstream therefrom, a socket at the upstream end of the flame tube forming with the burner body a spigoting connection engageable and disengageable in direction of flow, and retaining means to restrain said flame tube against involuntary displacement.
  • a combustion apparatus as claimed in claim 1, having means securing said burner body di: rectly to said structural member, said fuel supply passage and body having registering ports which are brought into sealed fluid-flow engagement when said body is secured on said structural member.
  • a combustion apparatus as claimed in claim 1, having means securing said burner body directly to the said structural membe:, said burner body having a base mounting provided with a port, said port registering with a complementary port formation of the supply passage and cooperating with said formation to form a spigot and socket joint, and said ports being brought into sealed fluid-flow engagement when said body is secured on said structural member.
  • a combustion apparatus as claimed in claim 1 having a supply connection projecting from said structural member, said burner body being directly and releasably connected thereto.
  • a combustion apparatus comprising an annular duct carrying a combustion-supporting air flow, a plurality of partitions attached at their ends to the inner and outer walls respectively of the duct and extending radially across the duct to divide it into segments, through each of which partitions passes a fuel supply passage, a like number of flame tubes annularly disposed within and extending longitudinally of said duct, each being located downstream of one of said partitions, a like number of fuel injection burners, each receiving fuel from one of said supply passages and being detachably mounted on the downstream side of a partition and having a body extending downstream therefrom, a socket at the upstream end of each flame tube forming with the burner body a spigoting connection engageable and disengageable in the direction of flow, and retaining means to restrain the flame tubes against involuntary displacement.
  • a combustion apparatus comprising an axial flow compressor having an annular axially directed outlet, an annular duct receiving a combustion supporting air flow from said compressor outlet, a plurality of partitions attached at their ends to the inner and outer walls respectively of the outlet, and extending radially across the outlet to divide it into segments, through each of which partitions passes a fuel supply pas-v sage.
  • a like number of flame tubes annularly disposed within and extending longitud nally of said duct. each flame tube being heat -er downstream of one of said partitions a like number of fuel injection burners.
  • each receiving fuel from one of said supply passages and being detachably mounted on the downstream side of a partition and having a body extending downstream therefrom, a socket at the upstream end of each flame tube forming with the burner body a spigoting connection engageable and disengageable in the direction of flow, and retaining means to restrain the flame tubes against involuntary dis placement.
  • a gas turbine plant comprising an axial flow compressor having an annular axially directed outlet, an annular duct receiving the output from said compressor outlet, a turbine nozzle annulus receiving the output from saidduct and an axial flow turbine, said compressor, duct, nozzle annulus and turbine being coaxially arranged for substantially straight through flow, a plurality of partitions attached at their ends to the inner and outer walls respectively of the outlet, and extending radially across the outlet to divide it into segments, a like number of flame tubes annularly disposed within and extending longitudinally of said duct, each flame tube being located downstream of one of said partitions, a like number of fuel injection burners, each being detachably mounted on the downstream side of a partition and having a body extending downstream therefrom, a socket at the upstream end of each flame tube forming with the burner body a spigoting connection engageable and disengageable in the direction of flow, and retaining means to restrain the flame tubes against involuntary dis lacement.
  • each partition element supporting a burner has a radiallv extending passa e therein and an axially directed aperture in its down stream wall which communicates with said radial passage, and wherein standing fuel supply ducting is accommodated in the passages thus formed, and the burner is mounted on said do nstream wall.
  • said burner having fuel receivin d ct n mating with said fuel connections to receive fuel therefrom.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Gas Burners (AREA)
  • Pressure-Spray And Ultrasonic-Wave- Spray Burners (AREA)
US641145A 1945-01-16 1946-01-14 Construction and mounting of fuel burner nozzles, especially for gas turbines and like apparatus Expired - Lifetime US2575264A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB1354/45A GB591113A (en) 1945-01-16 1945-01-16 Improvements relating to the construction and mounting of fuel burner nozzles, especially for gas turbines and like apparatus

Publications (1)

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US2575264A true US2575264A (en) 1951-11-13

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US641145A Expired - Lifetime US2575264A (en) 1945-01-16 1946-01-14 Construction and mounting of fuel burner nozzles, especially for gas turbines and like apparatus

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US (1) US2575264A (xx)
BE (1) BE462369A (xx)
CH (1) CH261477A (xx)
FR (2) FR920036A (xx)
GB (1) GB591113A (xx)
NL (1) NL70888C (xx)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2675675A (en) * 1954-04-20 Muctlpefi combustion chamber jet
US2690648A (en) * 1951-07-03 1954-10-05 Dowty Equipment Ltd Means for conducting the flow of liquid fuel for feeding burners of gas turbine engines
US2692478A (en) * 1951-02-24 1954-10-26 Boeing Co Turbine burner incorporating removable burner liner
US2697032A (en) * 1949-05-05 1954-12-14 Phillips Petroleum Co Acetylene gas producer
US2711631A (en) * 1949-06-21 1955-06-28 Hartford Nat Bank & Trust Co Gas turbine power plant
US3062482A (en) * 1952-08-25 1962-11-06 A V Roe Canada Ltd Gas turbine engined aircraft
US3741483A (en) * 1971-12-10 1973-06-26 Mitsubishi Heavy Ind Ltd Combustion air supply arrangement for gas turbines
EP0742412A1 (en) * 1995-05-12 1996-11-13 General Electric Company Dome assembly for a multiple annular combustor

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2670600A (en) * 1947-06-17 1954-03-02 Bristol Aeroplane Co Ltd Air distribution system for flame tubes of gas turbine engines
FR998079A (fr) * 1958-08-22 1952-01-14 Snecma Dispositif pour l'entrée de l'air dans la zone primaire d'une chambre de combustion de turbo-machine

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1466006A (en) * 1922-09-14 1923-08-28 Trood Samuel Apparatus for producing and utilizing combustible mixture
US1517422A (en) * 1922-08-05 1924-12-02 Leo G Hall Combination land and water vehicle
US1950046A (en) * 1933-05-12 1934-03-06 Surface Combustion Corp Gas burner
US1995417A (en) * 1933-05-12 1935-03-26 Surface Combustion Corp Gas burner
US2072731A (en) * 1934-12-03 1937-03-02 Steam Motors Inc Oil burner
US2224472A (en) * 1938-05-02 1940-12-10 Chandler Evans Corp Pressure fed carburetor
US2360130A (en) * 1941-03-26 1944-10-10 Armstrong Siddeley Motors Ltd High-speed propulsion plant
US2422214A (en) * 1944-08-17 1947-06-17 Westinghouse Electric Corp Fuel injector mounting for combustion chambers
US2432359A (en) * 1947-12-09 Internal-combustion turbine power
US2439273A (en) * 1948-04-06 Turbo-jet engine for aircraft

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2432359A (en) * 1947-12-09 Internal-combustion turbine power
US2439273A (en) * 1948-04-06 Turbo-jet engine for aircraft
US1517422A (en) * 1922-08-05 1924-12-02 Leo G Hall Combination land and water vehicle
US1466006A (en) * 1922-09-14 1923-08-28 Trood Samuel Apparatus for producing and utilizing combustible mixture
US1950046A (en) * 1933-05-12 1934-03-06 Surface Combustion Corp Gas burner
US1995417A (en) * 1933-05-12 1935-03-26 Surface Combustion Corp Gas burner
US2072731A (en) * 1934-12-03 1937-03-02 Steam Motors Inc Oil burner
US2224472A (en) * 1938-05-02 1940-12-10 Chandler Evans Corp Pressure fed carburetor
US2360130A (en) * 1941-03-26 1944-10-10 Armstrong Siddeley Motors Ltd High-speed propulsion plant
US2422214A (en) * 1944-08-17 1947-06-17 Westinghouse Electric Corp Fuel injector mounting for combustion chambers

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2675675A (en) * 1954-04-20 Muctlpefi combustion chamber jet
US2697032A (en) * 1949-05-05 1954-12-14 Phillips Petroleum Co Acetylene gas producer
US2711631A (en) * 1949-06-21 1955-06-28 Hartford Nat Bank & Trust Co Gas turbine power plant
US2692478A (en) * 1951-02-24 1954-10-26 Boeing Co Turbine burner incorporating removable burner liner
US2690648A (en) * 1951-07-03 1954-10-05 Dowty Equipment Ltd Means for conducting the flow of liquid fuel for feeding burners of gas turbine engines
US3062482A (en) * 1952-08-25 1962-11-06 A V Roe Canada Ltd Gas turbine engined aircraft
US3741483A (en) * 1971-12-10 1973-06-26 Mitsubishi Heavy Ind Ltd Combustion air supply arrangement for gas turbines
EP0742412A1 (en) * 1995-05-12 1996-11-13 General Electric Company Dome assembly for a multiple annular combustor

Also Published As

Publication number Publication date
CH261477A (de) 1949-05-15
NL70888C (xx)
GB591113A (en) 1947-08-07
FR920036A (fr) 1947-03-25
BE462369A (xx)
FR920038A (fr) 1947-03-25

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