Abstract: A plurality of turbulence promotors (11) are arranged on a surface of a stabilizer (11) comprising a stabilizer portion (14) contacting combustion gas and support portion (15) for rectifying flow except for a heat receiving surface (13) contacting with combustion gas to promote heat transfer from the stabilizer to an air fuel flow. The distance between the turbulence promotor closest to the heat receiving surface (13) and the heat receiving surface is at least 4 times the height of said turbulence promotor to prevent back fire due to provision of the turbulence promotor.

Abstract: A premix liquid fuel nozzle has longitudinal air entrance slots (24) into a cylindrical chamber (20). A centerbody (42) produces an axially increasing flow area toward the chamber outlet (32). Liquid fuel is atomized in a specified location (58) adjacent the conical centerbody (42). This area has a high axial shear velocity producing thorough vaporization and uniform mixing before combustion.

Abstract: Lean direct injection is used in a gas turbine combustor to reduce NO.sub.x emissions. The combustor has a multi-hole perforated plate disposed at the head end of the combustion chamber. The plate has a plurality of independent fuel jets and air jets for separately injecting fuel and air, respectively, into the combustion chamber in amounts which produce a lean fuel-air equivalence ratio. The perforated plate permits the fuel and air to be injected separately without any mixing outside of the combustion chamber. Fuel is delivered to the fuel jets by a fuel manifold having a branch connected to each one of the fuel jets. Air can be delivered to the air jets by an air plenum chamber located upstream of the perforated plate and in fluid communication with the air jets.

Abstract: A fuel/oxidizer pre-mixing combustion chamber for a turbojet engine is disclosed in which the pre-mixing device is incorporated into the upstream end wall structure of the combustion chamber. The upstream end portion of the combustion chamber, which is in communication with the oxidizer, has a plurality of generally radially extending, "V"-shaped members oriented such that the vertex of the "V" shape faces in an upstream direction. These members extend generally radially between the inner and outer walls which define the boundaries of the combustion chamber. The members define upstream edges and downstream edges which are axially the upstream edges. Fuel injectors extend radially inwardly between these members and are axially positioned between the locations of the upstream and downstream edges so as to spray fuel onto the adjacent sides of each of the members.

Abstract: In a combustor including a premixing type combustion burner which has an atomizer for ejecting a liquid fuel together with combustion air to atomize the liquid fuel, the atomizer is comprised of an inner shell to an inner peripheral surface of which the liquid fuel is supplied, an outer shell defining a passage for the combustion air running substantially straightly between the outer shell itself and an outer peripheral surface of the inner shell, and a swirling-flow guide plate for swirling the combustion air passed into the inner shell, while directing it in a downstream direction. The combustor further includes a resistor abruptly decreased in sectional area downstream and provided substantially downstream of the center of the swirling flow and in the vicinity of an outlet of the premixing type combustion burner for providing a resistance to a premixture ejected from the premixing type combustion burner.

Abstract: A gas turbine combustor for a gas turbine power plant comprises a combustion liner connected to a turbine and provided with a main fuel nozzle assembly and a sub-fuel nozzle assembly for jetting fuel to an inside of the combustion liner through nozzle holes, a base fuel supply line, a main fuel line for supplying a fuel to the main nozzle assembly for premixing an air with the fuel jetted through the nozzle hole for carrying out a lean burning in the combustion liner, and a plurality of sub-fuel lines for supplying the fuel to the sub-fuel nozzle assembly for mixing the fuel with a combustion air for carrying out a diffusion burning in the combustion liner. The main and sub-fuel lines branch off a front end of the base fuel supply line.

Abstract: A premix gas nozzle has longitudinal tangential entrance slots to a cylindrical chamber. There is an axially increasing flow area toward the chamber outlet, with pilot fuel centrally introduced near the outlet. A lean mix low NOx fuel nozzle is thereby stabilized.

Abstract: In order to enhance fuel injection efficiency, and particularly to atomize low fuel flows at high velocity with low fuel pressure, a fuel injection system 24 for a combustor 22 of a turbine engine 20 has an air blast tube 48 and a fuel supply tube 76. The air blast tube 48 is mounted at an acute angle relative to a wall 26 of the combustor 22 and has a first end 50 in communication with the combustor 22 and a second, enlarged end 52 disposed at an acute angle to an axis 54 of the air blast tube 48 in communication with a source of compressed air externally of the wall 26 of the combustor 22. The air blast tube 48 is operable to deliver compressed air from the source into the combustor 22.

Abstract: A combustor includes radially spaced apart outer and inner liners joined together at upstream ends through an annular dome disposed coaxially about a longitudinal centerline axis. A plurality of circumferentially spaced apart fuel injectors are disposed adjacent to an arcuate a splashplate for injecting fuel thereagainst, the splashplate being arcuate about the centerline axis. Primary air is channeled along the splashplate for mixing with the injected fuel for forming a fuel/air mixture dischargeable from the splashplate into a combustion zone defined between the liners and the dome.

Abstract: Improved fuel atomization for turbine engines utilizing high viscosity fuels is accomplished in an engine having an annular combustor (26) by providing an air injector (52) which produces an air stream (50) generally tangentially directed into the combustor (26) to intersect a fuel spray or film (46) also generally tangentially directed into the combustor (26) by a fuel injector (48) to atomize fuel for ignition by an igniter (60) in an annular outer wall (34) of the combustor (26).

Abstract: A combustion chamber suitable for use in a gas turbine engine includes a first enclosure with a fuel injector for low-power operation and a primary-oxidizer intake, a second enclosure with a fuel injector for full-power operation and a primary-oxidizer intake, and a third enclosure from which to evacuate burnt gases and which communicates with the first and second enclosures. The wall of the first enclosure includes intake orifices for a dilution oxidizer, but the wall bounding the second enclosure lacks any orifices other than the primary-oxidizer intake orifices.

Abstract: Difficulties with fuel atomization at low fuel flow rates in gas turbine fuel injection systems may be overcome by use of an impingement injector (52) including a cylindrical orifice (64, 94, 114) that is smooth, straight and uninterrupted by burrs or other disturbances having an exit opening (68, 96, 118) transverse to the axis of the orifice (64, 94, 114) and facing an impingement surface (72, 101, 126) that is planar and spaced from the exit opening (68, 96, 118).

Abstract: In order to efficiently atomize a fuel/air mixture to reduce smoke, a fuel injection system (26) for a combustor (12) of a turbine engine (10) includes an air blast tube (16), a fuel supply passage (42), and an impingement surface (30). The air blast tube (16) is mounted in relation to an opening (18) in a wall (20) of the combustor (12) such that a first end (22) of the tube (16) is in communication with the combustor (12) and a second end (24) of the tube (16) is in communication with a source of compressed air externally of the combustor (12) to facilitate the delivery of compressed air from the source into the combustor (12).

Abstract: The high cost of fuel injection nozzles for a gas turbine engine may be reduced by utilizing nozzles (60) formed of simple tubes (63) of capillary cross section having an integral flange (62) formed on one end (86) thereof.

Abstract: Complexity in a fuel injection system for a gas turbine engine including an annular combustor (26) may be reduced by utilizing a plurality of circumferentially spaced fuel injectors (46) within the annular combustion (26). First and second conduit branches (108, 110; 120, 122) in fluid parallel with one another are provided with one of the branches (108; 120) being connected to at least one of the injectors (46-1; 46-1, 46-3, 46-5) and the other of the branches (110; 122) being connected to the remainder of the injectors (46-2 through 46-6, inclusive; 46-2, 46-4, 46-6). A pump (102) interconnects a source of fuel (100) and the branches (108, 110; 120, 122). A normally closed, pressure responsive valve (114) isolates the other branch (110, 122) from the source (100) at low pressures typical of the starting of the turbine engine and is openable at somewhat higher pressures to allow fuel to flow to all of the injectors (46).

Abstract: A propulsion system comprises a gas turbine engine having sructurally integrated support systems that maximize structural efficiency and minimize cost.

Abstract: Poor atomization of fuel at low fuel flows in a gas turbine engine may be overcome in an engine including a rotary compressor (14), a turbine wheel (22) coupled to the compressor (14) to drive the same and a nozzle of (26) for directing gases of combustion at the turbine wheel (22). An annular combustor (34) defines a combustion annulus about the turbine wheel (22) and has an outlet (30) connected to the nozzle (26). A plurality of air blast tubes (50) include exit openings (52) circumferentially spaced about the combustor (32) and each has an axis (54) that is generally circumferentially directed with respect to the combustion annulus. A fuel discharge orifice (64) is located within each air blast tube (50) and a flange (66) is aligned with each orifice (64) in the path of fuel being discharged therefrom and at a substantial acute angle (.theta.) to the axis (54).

Abstract: A gas turbine engine fuel burner suitable for burning diesel fuel oil comprises an annular body and a center body coaxially located within the annular body so that an annular flow path is defined between them. At the upstream end of the burner an annular fuel manifold cooperates with the annular body and the center body to define two coaxial passages which direct air into the annular flow path with minimal turbulence. The low level of turbulence within the fuel burner reduces the possibility of spontaneous combustion occurring within the burner. The center body is hollow and is provided at its downstream end with an end cap which cooperates with the downstream end of the annular body to define a radial fuel/air mixture outlet.

Abstract: The expense of fabricating an annular combustor (10) for a gas turbine is minimized by providing a combustor housing (12) including an axially extending sleeve (14) and an annular liner (18) disposed within the housing (12) and about the sleeve (14). The annular liner (18) has concentric inner and outer axially elongated walls (20, 22) spaced from the sleeve (14) and the housing (12), respectively, and also has a radially extending wall (24) spaced from the housing (12) and interconnecting the inner and outer walls (20, 22) at one end to define a combustion chamber (26). The liner (18) to spaced from the housing (12) the sleeve (14) to define a compressed air flow path (28) extending from a radially outer compressed air inlet (30) to a radially inner compressed air outlet (32) in communication with the combustion chamber (26) axially remote from the radially extending wall (24).

Abstract: In order to reduce weight while also minimizing volume and pressure drop in a combustor assembly (18) for a turbine engine, the combustor assembly (18) includes an annular combustion chamber having a combustion space (20) defined entirely be an annular combustor case (22). The combustion chamber (18) is adapted to combust fuel and air in the combustion space (20) to generate gases of combustion. It has a combustor outlet (26) leading to a turbine nozzle (16). The outlet (26) is defined by a pair of turbine shrouds (28, 30). It also has a combustor inlet (32) leading to the combustion space (20). The combustor inlet (32) is defined by the combustor case (22) and one of the turbine shrouds (28). The combustion chamber (18) is adapted to receive injected air at the combustor inlet (32) in a manner creating a generally annular air flow in the combustion chamber (18) about the combustor case (22).

Abstract: Manifold head effects at low fuel flows in a fuel injected air breathing turbine are minimized by utilizing fuel injectors having fuel injecting tubes (66) with open ends (70) for fuel injection and provided with elongated capillary tubes (88) upstream thereof and connected to receive fuel from a fuel manifold (48) while uniform, relatively low velocity fuel exit flow from the ends (70) the injecting tubes (66) is achieved through the use of internal impingement surfaces (96, 102, 106, 110, 124).

Abstract: Improved fuel atomization for turbine engines operating at low fuel flows and at high altitudes is accomplished in an engine having an annular combustor 26 by utilizing start fuel injectors 46 provided with discharge orifices 56 and planar impingement surfaces 60 within the path of fuel discharged from the orifices 56 to provide flat sprays 62 that are generally tangential to the annular combustor 26 so as to be intercepted by an igniter 70 in a radial wall 34 of the combustor 26.

Abstract: An improved fuel delivery system for the combustor of a gas turbine engine is disclosed. The improvement comprises a combustor support pin, a fuel atomizer, and a tube for delivering fuel to the atomizer. The combustor support pin provides axial support for the combustor and includes a duct therethrough. The fuel delivery tube is removably located within the duct.

Abstract: Improved fuel atomization for turbine engines operating at low fuel flows and at high altitudes is accomplished in an engine having an annular combustor 26 by utilizing fuel injectors 46 provided with radially inward discharge orifices 56 and planar impingement surfaces 60 within the path of fuel discharged from the orifices 56 to provide a flat spray 62 that is generally tangential to the annular combustor 26.

Abstract: A carbureting device for the combustor of a gas turbine comprises an annular spin chamber which is defined by a generally annular housing having a closed forward end, an exhaust tube which is partially disposed within the spin chamber, and a flange disposed around the exhaust tube. A swirling air entrance is disposed between the annular housing and the flange and a fuel entrance is disposed along the annular housing spaced axially forward of the swirling air entrance. In one embodiment of the invention the exhaust tube is a venturi. In another embodiment a fairing is disposed outside and aft of the spin chamber and is sized and disposed about and extending at least to the end of the venturi, thereby forming a passageway between the fairing and the venturi. A second air swirling means is disposed in the passageway formed by the fairing and the venturi.

Abstract: A spray nozzle including a body having a fluid inlet at one end thereof and an exit orifice at the other end thereof. The fluid inlet communicates with the exit orifice through a fluid flow path extending through the body. The exit orifice is adapted to atomize the fluid as the fluid leaves the body. The spray nozzle also includes a separate portion for atomizing the fluid independent of the exit orifice. The separate portion communicates with the fluid flow path upstream of the exit orifice. The separate portion produces a flat spray of the fluid as the fluid leaves the body. With these features, the spray nozzle is well suited as a fuel injector for a gas turbine.

Abstract: Excellent fuel atomization in a turbine engine may be obtained with fuel injectors 46 including elongated, laminar discharge orifices 72 and impingement surfaces 76 disposed in the path of fuel 78 being discharged through the discharge orifices 72. Preferably, the fuel 78 is discharged as a flat spray generally tangentially to the annular combustion space 40 of an annular combustor 26.

Abstract: Excellent fuel atomization in a turbine engine may be obtained with fuel injectors 46, 47 including elongated, laminar discharge orifices 72, 100 and impingement surfaces 76, 102 disposed in the path of fuel 78, 106 being discharged through the discharge orifices 72, 100. Preferably, the fuel 78, 106 is discharged as a flat spray generally tangentially to the annular combustion space 40 of an annular combustor 26.

Abstract: The expense of fabricating an annular combustor for use with a small diameter gas turbine engine is minimized in a construction including a bell-shaped housing 52, an axially extending sleeve 66 within the housing 52, an annular liner 74 within the housing 52 and about the sleeve 66 and including inner, outer and radial walls 80, 76, 78, respectively, spaced from the sleeve 66 and the housing 52 and all being formed of sheet metal. A circular fuel manifold 108 with angularly spaced fuel dispensing openings 122 is disposed between the housing 52 and the outer wall 76 and a plurality of open ended, elongated tubes 112 extend through the outer wall 76 and are in fluid communication with respective openings 122 in the fuel manifold 108 to introduce air and fuel generally tangentially into the liner 74.

Abstract: Combustion apparatus for a gas turbine engine comprises a burner which is so configured and located within a combustion chamber so as to urge fuel and air mixture ejected therefrom into a fuel rich toroidal vortex in an upstream first combustion zone of the combustion chamber. Unburnt fuel from the first combustion zone is mixed with additional air in a second fuel weak combustion zone downstream of the first zone. Adjustment of the air to fuel ratios in the two combustion zones results in the reduction of smoke and oxides of nitrogen reduction.

Abstract: An improved air swirler and fuel injector for a gas turbine combustor system employs primary fuel injection into a primary recirculation combustion zone supplied by an inner annulus of swirling air at high fuel-air ratio under low power conditions. A selectively staged secondary fuel flow is supplied into an outer annulus downstream of secondary air swirlers supplying a secondary recirculation zone for combustion at higher fuel flows but lower fuel-air ratios under high power conditions. The secondary fuel flow may be separately injected into the outer annulus through conduit means extending through the primary air swirler vanes. Alternatively, the secondary fuel flow may be combined with the primary fuel in the fuel injector body and the momentum due to mass flow used to inject the fuel into the outer annulus.

Abstract: In the dual burner of a gas turbine or a hot gas generating system, the swirler (1) is formed from at least two doubly curved sheets (4, 5) subject to tangential air inlet (7). These sheets (4, 5) are folded along diagonals (10a, 10b) running conically outwards in the outlet flow direction. One curved fold face forms an inner cone (4b, 5b) expanding in the direction of the swirler outlet end while the other curved fold faced forms an outer cone (4a, 5a) which contracts in the direction of the end outlet. The inner cones (4b, 5b) each carry at their ends a fuel main (8) whose fuel nozzles (9) are directed radially inwards towards the inner zone of the swirler (1). The liquid fuel (2a, 2b) is directed onto the outer cone (4a, 5a), the oil film (6) forming there being "rolled in" by the air (7) flowing into the outer cone (4a, 5a).

Abstract: A dual fuel injector for a gas turbine power plant has fuel ducting including gas and liquid fuel ducts which can be removed easily for access to a nozzle of the fuel duct. A separate fluid ducting assembly is attached in the head of each combustor can and has an annular duct, a passage and drillings. The power plant is started on liquid fuel, and runs subsequently on gas produced in a coal gasifier. The flow of air through the duct and passage prevents any unburnt fuel or combustion products from entering the gas duct, and the flow through the drillings prevents carbon build up on face of the injector.

Abstract: A connection device is disclosed for connecting a burner ring or flame holder made of a composite material and having a "V"-shaped cross-section, to an afterburner chamber duct of a turbojet engine. The connecting device has a cylindrical portion and a trapezoidal portion. The cylindrical portion extends through hollow projections formed from the composite ring material and extending from the ring in an upstream direction. The converging, opposite sides of the trapezoidal portion bear against inner surfaces of the converging flanges forming the "V" cross-section of the ring. The cylindrical portion is then attached to the afterburner duct via a link-bar mechanism.

Abstract: Combustors [22] which employ such features as rotary cup fuel atomizers and the latter in conjuction with CIVIC combustion to enhance flame performance, to produce efficient combustion over a wide operating range, and to eliminate the problems typically encountered in liquid fueled combustors, especially as they are scaled down in size. A porous or perforate barrier [180] can be used to keep large fuel drops from reaching the combustion zone [124] of the combustor. Also, an air buffer system can be employed to keep fuel from backing up to the bearings [32] by which the rotary cup [98] of the atomizer is rotatably supported, and a fuel supply line [56] with a strategically located orifice [176] can be provided to effect an interaction between fuel discharged therefrom and the buffer air. This keeps the film of fuel formed on the atomizer cup [98] from being disrupted and also keeps the buffer air from penetrating to an inner hot gas recirculation zone [130] in the combustor.

Abstract: A ring burner of a ring combustion chamber is divided into a large number of honeycomb-like parallel axis canals for the combustion air, by radial and circumferential plate canals or by radial plate canals, longitudinal tubing, radial tubing and annular tubing into which combustion gas is introduced from nozzles in the surrounding walls. At the burner outlet there are flame retention nozzles provided above the frontal surface area of the plate canals or tubes. Fuel nozzles are provided in front of the burner inlet for operation as a dual burner with gaseous and liquid fuels.

Abstract: An improved prechamber and fuel manifold structure for a gas turbine engine having a premixing-prevaporizing type combustor, the improvement residing in the provision of a prechamber having an internal cylindrical surface swept by swirling pressurized air passing through the prechamber from a plenum to the combustor and in the provision of a fuel manifold having a plurality of individual fuel conduits extending from a remote fuel supply pipe to a corresponding plurality of delivery heads in the prechamber. The delivery heads direct fuel generally tangent to the internal cylindrical surface to form a film for vaporization in the passing air and the length of the fuel conduits and the flow areas of the fuel conduits and the delivery heads are generally equal so that fuel flow in the fuel conduits is balanced and so that when the conduits are opened to atmospheric pressure at termination of combustion, purging of residual fuel to foreclose coking is rapid and complete.

Abstract: A gas turbine fuel injection nozzle is disclosed in which the secondary fuel is spread into a very thin film entirely within a region of low air momentum. The fuel is therefore not affected by turbulence, and this results in an evenly circumferentially distributed fuel film at the discharge orifice of the nozzle resulting in an even and extremely fine spray of fuel to enhance proper engine performance.

Abstract: Combustion systems which employ such features as rotary cup fuel atomizers and the latter in conjunction with CIVIC combustion to enhance flame performance, to produce efficient combustion over a wide operating range, and to eliminate the problems typically encountered in liquid fueled combustors, especially as they are scaled down in size. Innovations that can advantageously be employed in combustors operated on other fuels are also disclosed.

Abstract: In order to prevent the accumulation of carbon on the pintle of a gas turbine engine fuel burner, the pintle has a hollow base and inlet holes so that a small mass flow of compressor delivery air can enter the hollow base and flow over its interior surface.

Abstract: A combustor assembly for an automotive gas turbine engine includes a continuously ignited, pilot flame tube supported on a combustor dome immediately downstream of a prevaporization prechamber with a fuel and air swirler. The pilot flame tube has an air swirler and fuel nozzle supported on an interior bulkhead and is associated with a fuel system that maintains a shielded pilot flame to prevent flame-out of flammable air/fuel mixtures in the main reaction chamber of the combustor; the fuel system supplies greater fuel to the pilot flame tube at engine fuel start and stop to produce an extended flame plume that extends into the reaction chamber to burn residuum of fuel when the air/fuel ratio in the main reaction chamber mixture is below the flammability limit of the fuel.

Abstract: A combustion apparatus for gas turbine engines particularly adapted to reduce emissions to meet automotive requirements. The fuel is laid on the wall of a cylindrical prechamber and evaporated from the wall by combustion air which is introduced through a swirler at the upstream end of the prechamber. The inner surface of the prechamber is artificially roughened by a grid of grooves to improve fuel evaporation. The fuel is laid on the wall from an annular manifold extending around the upstream end of the prechamber through tangential orifices leading from the manifold into the interior of the prechamber. More air enters through entrance ports distributed around the prechamber toward its downstream end. The resulting lean fuel-air mixture is delivered past an annular flow dam at the outlet of the prechamber into a domed combustor wall forming a combustion or reaction zone which is abruptly enlarged from the prechamber.

Abstract: Fuel is injected through very fine holes normally into a swirled airstream. A finally atomized fuel results and a solid cone of fuel and air passes into the combustion chamber. The injector gives high combustion efficiency at idle conditions reduces smoke at high power conditions.

Abstract: The disclosure of this invention pertains to an annular combustion chamber for a gas turbine engine. The chamber has at its upstream end an annular array of baffles defining inlet passages for air or air/fuel mixture. The baffles overlap in the circumferential sense so that air or air/fuel mixture entering passages is directed by the baffles into a circumferential swirl. Those surfaces of the baffles which are open to the interior of the chamber are swept by the fluid emerging from the passages to be cooled thereby.

Abstract: A potted head combustion chamber is provided with an annular gap between the downstream end of the pot and the combustion chamber itself. Air is supplied radially inwardly through this gap to remove any fuel which adheres to the walls of the pot, and prevents it from reaching the walls of the combustion chamber. Fuel on the combustion chamber walls can cause localized burning of the fuel with a consequent increase in temperature of that part of the combustion chamber wall.